diff --git "a/logs/app.log" "b/logs/app.log" new file mode 100644--- /dev/null +++ "b/logs/app.log" @@ -0,0 +1,63294 @@ +2025-04-03 at 08:50:36 | INFO | src.config:update_log_path:253 - Additional logs will be stored in: /workspace/DeepSearch/trainer_output_meta-llama_Llama-3.2-3B-Instruct_gpu0_20250403_085036/logs +2025-04-03 at 08:50:36 | INFO | __main__::33 - Training output directory: /workspace/DeepSearch/trainer_output_meta-llama_Llama-3.2-3B-Instruct_gpu0_20250403_085036 +2025-04-03 at 08:50:36 | INFO | __main__::34 - Logs are being saved to both ./logs and /workspace/DeepSearch/trainer_output_meta-llama_Llama-3.2-3B-Instruct_gpu0_20250403_085036/logs +2025-04-03 at 08:50:36 | INFO | __main__::37 - Initializing model meta-llama/Llama-3.2-3B-Instruct +2025-04-03 at 08:51:21 | INFO | __main__::48 - Setting up LoRA adapter +2025-04-03 at 08:51:27 | INFO | __main__::59 - Loading datasets +2025-04-03 at 08:51:27 | INFO | __main__::61 - Loaded 286 training examples and 32 test examples +2025-04-03 at 08:51:27 | INFO | __main__::64 - Setting up training arguments +2025-04-03 at 08:51:27 | INFO | __main__::100 - Setting up verifier +2025-04-03 at 08:51:27 | INFO | __main__::112 - Initializing trainer +2025-04-03 at 08:51:28 | INFO | __main__::131 - Starting training +2025-04-03 at 08:51:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:51:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inclinatoin angle Earth orbit plane equatorial +2025-04-03 at 08:51:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "orbital inclination of Earth or Titian equatorial intersection" +2025-04-03 at 08:51:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the acute angle formed at the intersection of the orbit plane and the equatorial plane of a specific celestial body, such as a planet or star? +2025-04-03 at 08:51:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: angle of inclination of orbit vs equatorial plane of a celestial body +2025-04-03 at 08:51:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:51:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the inclination of a celestial body's orbit, i.e., the acute angle formed at the intersection of the orbit plane and the body's equatorial plane, in degrees? +2025-04-03 at 08:51:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: incline of orbit plane at intersection with earth equatorial plane +2025-04-03 at 08:51:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 08:51:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:51:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the definition and notation of "inclination" in aerospace engineering, specifically the acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, in degrees? +2025-04-03 at 08:51:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 08:51:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inclination angle at intersection with earth equatorial plane for high inclination ordinate orbit +2025-04-03 at 08:51:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:51:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Definition of inclination angle measurements in space mission planning and orbit determination +2025-04-03 at 08:51:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 08:51:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: star-angle difference and inclination angle +2025-04-03 at 08:51:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:51:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard unit of measurement for the inclination angle formed at the intersection of the orbit plane and the reference body's equatorial plane? +2025-04-03 at 08:51:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 08:51:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: star-angle difference of 1.12 degrees and spacecraft inclination angle +2025-04-03 at 08:51:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 08:51:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:51:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:51:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:51:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:62 - Student lengths: [889, 328, 363, 367, 508, 509] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:64 - Average student length: 494.00 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 11.00 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_correctness:66 - Length ratio: 44.91 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.417 ± 0.398 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 2.11 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 2, 5, 0, 0, 5] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:51:52 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:51:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:52:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was added to the Apollo lunar sample return container +2025-04-03 at 08:52:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 08:52:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar sample return tote bag contents +2025-04-03 at 08:52:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:52:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:52:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo suit vacuum brush tote bag contents +2025-04-03 at 08:52:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 08:52:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar sample tote bag functionality and materials +2025-04-03 at 08:52:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 08:52:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Charged Particle Lunar Environment Experiment components and packaging +2025-04-03 at 08:52:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo experiment package materials and construction +2025-04-03 at 08:52:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo experiment package materials and construction (lunar surface deployment) +2025-04-03 at 08:52:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 08:52:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo experiment package materials and construction (heating and electronic components) +2025-04-03 at 08:52:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo electronics and sensor components (spring and flexible materials) +2025-04-03 at 08:52:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flexible spring material used in Apollo experiment connectors +2025-04-03 at 08:52:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:52:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: wire springs used in Apollo space missions +2025-04-03 at 08:52:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 08:52:17 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:52:17 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:52:17 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, True, True] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:62 - Student lengths: [756, 230, 573, 1377, 390, 603] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:64 - Average student length: 654.83 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 7.00 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_correctness:66 - Length ratio: 93.55 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.352 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 3.67 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 10, 0, 0] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 08:52:17 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...'] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 08:52:17 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:52:17 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:52:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Balance Control System operational mode during Fukushima Daiichi nuclear accident +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: system mode fluctuations observed in simultaneous illumination and temperature +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "system mode when observed fluctuations" +2025-04-03 at 08:52:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:52:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 recovery operations control system +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 optics shaft angle fluctuations zero optics mode +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "launch vehicle launch dynamics water quantity fluctuations" +2025-04-03 at 08:52:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-03 at 08:52:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 abort mode control system +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode cause and characteristics +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "launch vehicle exhaust impact on water quantity measurements" +2025-04-03 at 08:52:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 08:52:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 navigation and guidance system mode +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode resolver shaft motion effects and ventilation +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "water system mode contamination bacteria detection" +2025-04-03 at 08:52:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 08:52:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 guidance, navigation, and control system transearth phase +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode resolver slip-ring mechanism vacuum susceptibility +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "system mode for water quantity measurement spacecraft" +2025-04-03 at 08:52:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:52:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 transearth navigation system +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode vacuum susceptibility 5 ohms +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "system mode potable water quantity measurement anomaly 14.l.8" +2025-04-03 at 08:52:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:52:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode 5 ohms 0.3 degree degradation +2025-04-03 at 08:52:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "system mode potable water tank dormant procedure" +2025-04-03 at 08:52:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:52:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:52:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo zero optics mode Appolo 41 hours +2025-04-03 at 08:52:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 08:52:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "dormant system mode for potable water tank" +2025-04-03 at 08:52:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:52:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:52:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 sextant shaft random motion zero optics mode +2025-04-03 at 08:52:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 08:52:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "shutdown/fill mode water quantity measurement placeholder" +2025-04-03 at 08:52:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-03 at 08:52:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:52:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:52:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, False, True] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:62 - Student lengths: [28, 395, 28, 1020, 1917, 1495] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:64 - Average student length: 813.83 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_correctness:66 - Length ratio: 50.86 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.475 ± 0.475 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.00 ± 4.12 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 6, 9, 9] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 08:52:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-03 at 08:52:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe potable water quantity measurement fluctuated briefly on two occasions du...', 'Result 1:\nAt site 6, the record was similar to that for site 7 with an initial positive...', 'Result 1:\nAt site 6, the record was similar to that for site 7 with an initial positive...', 'Result 1:\nPreflight testing of both command module and lunar module water supplies reve...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nPotable water was obtained by periodically pressurizing the potable tank with...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...'] +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:52:46 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:52:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: data from Hafele-Keating experiment +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment to measure telluric current consisted of an electrode placed close to the launch site and two electrodes spaced approximately 2500 feet from the base electrode at a 90-degree included angle (shown in figure ll.l-2). The telluric current system failed to detect any launch effects. It was expected that the current would show an increase until the vehicle exhaust plume broke effective electrical contact with ground. The high density of metallic conductors in the ground near the launch site may have functioned as a short circuit, which would have negated the detection of any changes in the current level. + +11.1.4 Measurement of the Air/Earth Current Density +------ + +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "optics shaft angle fluctuation measurement amplitude" +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 trajectory data +2025-04-03 at 08:52:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 08:52:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:52:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lasser interferometer angle fluctuations Apollo +2025-04-03 at 08:52:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 08:52:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: amplitude of fluctuations in Voyager 1 optics shaft angle maneuver +2025-04-03 at 08:52:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:52:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:53:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: amplitude of 0.3 degree fluctuations in optics shaft angle Voyager 1 +2025-04-03 at 08:53:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 08:53:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:53:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 optics shaft servo loop +2025-04-03 at 08:53:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 08:53:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:53:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:53:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:53:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, False, True] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:62 - Student lengths: [516, 703, 839, 696, 1473, 562] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:64 - Average student length: 798.17 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_correctness:66 - Length ratio: 79.82 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.221 ± 0.226 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.50 ± 2.14 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 2, 1, 0, 6] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment to measure telluric current consisted of an electrode placed close to the launch site and two electrodes spaced approximately 2500 feet from the base electrode at a 90-degree included angle (shown in figure ll.l-2). The telluric current system failed to detect any launch effects. It was expected that the current would show an increase until the vehicle exhaust plume broke effective electrical contact with ground. The high density of metallic conductors in the ground near the launch site may have functioned as a short circuit, which would have negated the detection of any changes in the current level. + +11.1.4 Measurement of the Air/Earth Current Density +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 08:53:04 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 08:53:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: insulation failure in heating systems +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what causes tank insulation anomalies +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: causes of reduced heat loss in industrial storage tank insulation +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "tank-insulation anomalies causes" +2025-04-03 at 08:53:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: contaminant introduction tank helium loading annealing effect +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: causes of tank insulation failure Apollo 204 mission +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (helium storage tank explosion) + (catastrophic failure + vacuum insulation degradation) +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 electrical system failure causes correlation with insulation failure" +2025-04-03 at 08:53:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 08:53:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cold-soak thermal contamination helium tank +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 204 fuel cell oxygen tank insulation design changes +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (titanium hydrogen contamination) + vacuum insulation failure + tank explosion +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11-14 electrical system transformer issues" +2025-04-03 at 08:53:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:53:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: condensation-driven pressure rise tank cooling system +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 204 cryogenic oxygen tank electrical instrument safety +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (hydrogen contamination effects on vacuum insulation) + helium loading + heat exchanger lines +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module ascent power noise, pressure debris correlation" +2025-04-03 at 08:53:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:53:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: short-circuit fan circuit oxygen tank +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 204 anomaly Fuel Cell 2 Off loading +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (helium loading + hydrogen contamination + tank insulation degradation + temperature effects) +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 moon suit spacewalk contamination cause link" +2025-04-03 at 08:53:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:53:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical fault cryogenic oxygen tank +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 204 Command Module Ac bus B undervoltage issue +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (stainless steel vs aluminum + fire hazard prevention + cryogenic tank maintenance) +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar landing heat and pressure system failure effects" +2025-04-03 at 08:53:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 08:53:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell electrical fault cryogenic oxygen tank +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 204 lunar module mechanical noise and electrical issues +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (cryogenic oxygen tank plumbing + heat cycling in detanking + tank heater troubleshooting) +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "hydrogen tank pressure anomalies heated fuel system temperature relationship" +2025-04-03 at 08:53:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mechanical stress bay panel failure oxygen tank fuel cell failure +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (cryogenic oxygen tank design modifications + stainless steel sheathing + plumbing redesign) +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo fuel cell failure oxygen tank pressure emergency procedures" +2025-04-03 at 08:53:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:53:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank failure oxygen supply system shut down +2025-04-03 at 08:53:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 08:53:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (cryogenic oxygen tank design optimization + stainless steel sheathing + oxygen distribution system modifications + power conservation solutions) +2025-04-03 at 08:53:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:41 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:53:41 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:53:41 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1421, 1865, 1193, 1565, 1799, 478] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [27, 27, 27, 27, 27, 27] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:64 - Average student length: 1386.83 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 27.00 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_correctness:66 - Length ratio: 51.36 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.633 ± 0.448 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.50 ± 3.95 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [9, 7, 0, 9, 8, 0] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 08:53:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:53:41 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:53:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:53:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pilot astronaut +2025-04-03 at 08:53:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 08:53:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot Apollo 11 +2025-04-03 at 08:53:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:53:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:53:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pilot apollo mission +2025-04-03 at 08:53:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:53:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Michael Collins Apollo 11 +2025-04-03 at 08:53:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 08:53:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:53:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pilot apollo 13 +2025-04-03 at 08:53:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:53:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:53:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: commander apollo 13 +2025-04-03 at 08:53:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:53:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:53:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:53:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:53:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.17 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:62 - Student lengths: [174, 315, 325, 338, 228, 156] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:64 - Average student length: 256.00 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 20.00 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_correctness:66 - Length ratio: 12.80 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.318 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.00 ± 1.53 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 4, 0, 2, 0, 0] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe unprecedented powered-down state of the command module required generatio...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nMedical kits for future flights will include nose drops packaged the same as ...'] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 08:53:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:53:51 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:53:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:53:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Initial accelerometer value before translunar injection NASA Apollo" +2025-04-03 at 08:53:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:53:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:53:58 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:53:58 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:53:58 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, True] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:62 - Student lengths: [620, 516, 519, 1580, 597, 981] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:64 - Average student length: 802.17 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_correctness:66 - Length ratio: 133.69 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 0, 0, 0] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 08:53:58 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:53:58 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:53:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Jupiter IICExplore mission propellant capacity +2025-04-03 at 08:54:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 08:54:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, False, True] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:62 - Student lengths: [275, 805, 476, 33, 611, 557] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:64 - Average student length: 459.50 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_correctness:66 - Length ratio: 91.90 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 0, 0, 0, 0] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nAll service module reaction control parameters were normal from lift-off to t...'] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 08:54:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:54:04 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:54:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'materials for vehicle shades with UV resistance and strength-to-weight ratio' +2025-04-03 at 08:54:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 08:54:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'types of fabrics used for extravehicular visor assemblies or space suit shades ' +2025-04-03 at 08:54:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 08:54:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'material for Aclar shades or extravehicular visor assembly shades ' +2025-04-03 at 08:54:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 08:54:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:62 - Student lengths: [411, 692, 285, 883, 519, 636] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:64 - Average student length: 571.00 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_correctness:66 - Length ratio: 114.20 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.108 ± 0.242 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 1.12 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 3, 0] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 08:54:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:54:14 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:54:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "total oxygen supply after surge tank incident" +2025-04-03 at 08:54:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 08:54:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oxygen supply lunar module cargo bay" +2025-04-03 at 08:54:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 08:54:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oxygen supply surge tank lunar module" +2025-04-03 at 08:54:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:54:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "surge tank oxygen supply amount" +2025-04-03 at 08:54:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 08:54:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, True] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:62 - Student lengths: [326, 432, 347, 341, 28, 604] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:64 - Average student length: 346.33 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 11.00 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_correctness:66 - Length ratio: 31.48 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.079 ± 0.177 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 3.73 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 10, 0, 0, 0, 0] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 08:54:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:54:25 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:54:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude above oblate earth +2025-04-03 at 08:54:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 08:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the unit of measurement for the maximum altitude above the Earth's equator? +2025-04-03 at 08:54:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 08:54:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:54:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:62 - Student lengths: [674, 977, 479, 286, 108, 852] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:64 - Average student length: 562.67 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_correctness:66 - Length ratio: 140.67 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.117 ± 0.165 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.33 ± 0.47 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 1, 0, 0] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 08:54:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 08:54:32 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:54:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band communication duration system actuation +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the duration of nominal S-band communications from system actuation for a specific space mission or spacecraft system? +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nominal S-band communication duration during INS/GPS system actuation +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nominal s-band communication duration from system actuation for a specific space mission or vehicle +2025-04-03 at 08:54:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:54:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, True] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:62 - Student lengths: [656, 211, 706, 501, 582, 515] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:64 - Average student length: 528.50 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 22.00 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_correctness:66 - Length ratio: 24.02 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.233 ± 0.165 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 0.47 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 1, 1, 0, 0] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 08:54:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:54:39 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:54:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the type of insulation used in the S-Il thermal protection system? +2025-04-03 at 08:54:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:54:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:54:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:54:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:62 - Student lengths: [453, 436, 484, 365, 94, 946] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:64 - Average student length: 463.00 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_correctness:66 - Length ratio: 46.30 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 1, 0] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 08:54:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:54:44 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:54:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:54:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propellant mass after cryogenic ignition lunar mission +2025-04-03 at 08:54:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:54:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: product of inertia lunar module mass and distribution +2025-04-03 at 08:54:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:54:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:54:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hydrogen molecular weight liquid oxygen molecular weight then product of inertia lunar module descent propulsion +2025-04-03 at 08:54:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 08:54:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: product of inertia lunar module mass table A-1 +2025-04-03 at 08:54:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 08:54:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:54:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inertia temperature relationship +2025-04-03 at 08:54:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:54:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module product of inertia at ignition +2025-04-03 at 08:54:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:54:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:54:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module product of inertia after fueling +2025-04-03 at 08:54:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:54:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module moment of inertia after cryogenic propellant load +2025-04-03 at 08:55:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:55:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar probing and reaction control system propellant product of inertia +2025-04-03 at 08:55:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 08:55:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertia during module separation +2025-04-03 at 08:55:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:55:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:55:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:55:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, False, True, False] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1121, 633, 420, 1662, 544, 3282] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:64 - Average student length: 1277.00 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_correctness:66 - Length ratio: 255.40 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.242 ± 0.366 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.92 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 0, 0, 8, 0, 0] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 08:55:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nCommand module.- Command module reaction control system propellant usages can...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...'] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 08:55:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:55:04 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:55:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:55:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +scale factor std deviation X-axis accelerometer "accelerometer calibration error" +2025-04-03 at 08:55:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 08:55:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Standard deviation of X-axis accelerometer scale factor error ppm +2025-04-03 at 08:55:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 08:55:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:55:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:55:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, False, False] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:62 - Student lengths: [529, 614, 420, 570, 290, 574] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:64 - Average student length: 499.50 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_correctness:66 - Length ratio: 124.88 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.054 ± 0.121 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 1.12 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 0, 0, 0, 0, 0] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 08:55:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:55:12 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:55:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module in an aircraft or command and service module in an aircraft +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch and lift capacity +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of aircraft that can lift command module in space or orbital path +2025-04-03 at 08:55:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 08:55:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:55:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Skylab command module equivalent in aircraft +2025-04-03 at 08:55:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 08:55:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module lift-off data +2025-04-03 at 08:55:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:55:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module equivalent in military aircraft or heavy-lift aircraft +2025-04-03 at 08:55:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 08:55:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo trajectory lift-off data +2025-04-03 at 08:55:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 08:55:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spacecraft capable of carrying command module +2025-04-03 at 08:55:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 08:55:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module launch sequence +2025-04-03 at 08:55:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:55:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module technology and future space missions +2025-04-03 at 08:55:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:55:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module lunar descent data +2025-04-03 at 08:55:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:55:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo equivalent in spacecraft or technology for aircraft +2025-04-03 at 08:55:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 08:55:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V Command Module Lunar Module +2025-04-03 at 08:55:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:55:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module Command Module separation +2025-04-03 at 08:55:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:55:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module undocking procedures +2025-04-03 at 08:55:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:55:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Preventing gimbal lock during lunar module undocking +2025-04-03 at 08:55:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 08:55:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:55:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:55:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, True, True, True] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1042, 784, 414, 605, 1718, 508] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:64 - Average student length: 845.17 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 18.00 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_correctness:66 - Length ratio: 46.95 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.375 ± 0.425 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 3.54 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 0, 0, 0, 9, 1] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 08:55:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe performance of the comma...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe performance of the comma...', 'Result 1:\nUpon completion of photography, the two docked vehicles were maneuvered back ...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 08:55:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nDESCRIPTIONS·········· A-1 A.1 COMMAND AND ...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...'] +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:55:37 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:55:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:55:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the context of the state of propellant isolation valves during system decontamination in Hawaii? +2025-04-03 at 08:55:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 08:55:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:55:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:55:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, True, True] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:62 - Student lengths: [987, 990, 524, 604, 484, 477] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:64 - Average student length: 677.67 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 15.00 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_correctness:66 - Length ratio: 45.18 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 0, 0, 0, 0] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 08:55:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:55:44 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:55:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:55:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fit for simulating pressures like in high-altitude environments +2025-04-03 at 08:55:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 08:55:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure gauge discrepancy suit measurement diving +2025-04-03 at 08:55:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:55:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:55:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft suit pressure measurement discrepancy initial launch +2025-04-03 at 08:55:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:55:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electroless nickel plating particles cause of pressure transducer failure +2025-04-03 at 08:55:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:55:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo batch of nickel plating deposits 1960s equipment failure analysis +2025-04-03 at 08:55:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 08:55:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: failure analysis transformer isolation issue Apollo 12 +2025-04-03 at 08:55:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 08:55:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:55:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transducer wafer contamination repair procedures +2025-04-03 at 08:55:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 08:55:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:00 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:56:00 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:56:00 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, True, False] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:62 - Student lengths: [286, 646, 533, 508, 887, 272] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:64 - Average student length: 522.00 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 28.00 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_correctness:66 - Length ratio: 18.64 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.352 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.17 ± 2.19 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 0, 0, 6, 0] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 08:56:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nFigure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n...'] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 08:56:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:56:00 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:56:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:56:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo mission 11 lunar module undocking time" + +2025-04-03 at 08:56:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 08:56:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo 11 lunar module separation time from command module" + +2025-04-03 at 08:56:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 08:56:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo 11 lunar module separation time from service module" + +2025-04-03 at 08:56:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 08:56:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo 11 lunar module separation time from service module 4.75 hours before entry" + +2025-04-03 at 08:56:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 08:56:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo 11 lunar module separation time from service module 3 hours 28 minutes before service module separation" + +2025-04-03 at 08:56:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 08:56:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:56:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:56:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, True, True] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:62 - Student lengths: [384, 219, 417, 851, 481, 458] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:64 - Average student length: 468.33 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_correctness:66 - Length ratio: 78.06 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.158 ± 0.354 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.83 ± 1.86 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 5, 0, 0] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 08:56:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:56:14 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:56:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage of a Falcon Heavy LV0005, oxygen consumption per tank (Ascent stage Tank 2) +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage fuel and oxidizer consumption on Falcon 9 +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage Tank 2 oxygen consumption +2025-04-03 at 08:56:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 08:56:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:56:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage Tank 2 at LV0005, oxygen consumption rate +2025-04-03 at 08:56:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 08:56:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:56:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:56:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:62 - Student lengths: [152, 496, 551, 1199, 632, 510] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:64 - Average student length: 590.00 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 3.00 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_correctness:66 - Length ratio: 196.67 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.200 ± 0.206 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 0.75 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 2, 1, 1] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 08:56:27 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 08:56:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB velocity increase 19 hours 17 minutes +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vs unexplained velocity increase +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + + +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: track Apollo 11 spacecraft S-IVB unexplained velocity increase 19 hours 17 minutes +2025-04-03 at 08:56:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:56:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB velocity increase velocity 25665 ft/sec mixture ratio shifts +2025-04-03 at 08:56:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:56:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB deceleration and thrust calculation mixture ratio +2025-04-03 at 08:56:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 08:56:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:56:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB velocity increase 5 ft/sec duration 2l7 seconds +2025-04-03 at 08:56:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB engine cutoff 00:12:30 velocity increase +2025-04-03 at 08:56:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:56:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB engine cutoff initiation time 19 hours 00:00 +2025-04-03 at 08:56:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 08:56:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB translunar injection maneuver velocity increase +2025-04-03 at 08:56:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 08:56:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:56:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:56:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:56:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:62 - Student lengths: [715, 568, 456, 507, 873, 790] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:64 - Average student length: 651.50 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_correctness:66 - Length ratio: 81.44 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.358 ± 0.323 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.05 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 1, 0, 1, 0, 2] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 08:56:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 08:56:44 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 08:56:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:56:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system initial propellant consumption +2025-04-03 at 08:56:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 08:56:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system hydrazine propellant consumption initial consumption NASA spacecraft +2025-04-03 at 08:56:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 08:56:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:56:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:56:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:56:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, True, True, False] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:62 - Student lengths: [649, 455, 464, 574, 1170, 366] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:64 - Average student length: 613.00 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_correctness:66 - Length ratio: 61.30 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.100 ± 0.144 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 0.76 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 1, 2, 0] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 08:56:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 08:56:52 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:56:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "MEDICATION DROPS FOR TRANSPORTATION KITS" +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: air travel prescription eye drop packaging standards +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of packaging do eye drops for medical use and intended for use by passengers during a prolonged period on aircraft require? +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "types of drops packaged similarly to eye drops for in-flight medical kits" +2025-04-03 at 08:56:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:56:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:56:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal drops package design for medical kits on future flights +2025-04-03 at 08:56:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:56:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of packaging is used for eye drops in aircraft cabin pressurization systems or in emergency situations on commercial flights? +2025-04-03 at 08:56:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:56:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal spray bottles used as drops in medical kits +2025-04-03 at 08:56:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:56:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: standardized nasal and ocular drop packaging for space medical kits +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of eye drop packaging used in emergency oxygen mask or medical kits on airplanes +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal nose drops packaged like eye drops +2025-04-03 at 08:57:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:57:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: medical drop packaging for space applications +2025-04-03 at 08:57:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: international standards for packaging pharmaceutical eye drops for space +2025-04-03 at 08:57:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: FDA guidelines for packaging eye drops for medical use in space +2025-04-03 at 08:57:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 08:57:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: US space agency guidelines for medical emergency response kits +2025-04-03 at 08:57:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: International Organization for Standardization guidelines for medical kits on spacecraft +2025-04-03 at 08:57:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: medical kit standards for space missions +2025-04-03 at 08:57:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:13 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:57:13 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:57:13 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:62 - Student lengths: [589, 518, 735, 373, 423, 458] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:64 - Average student length: 516.00 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_correctness:66 - Length ratio: 51.60 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.433 ± 0.352 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 3.09 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 9, 0, 3, 3, 0] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 08:57:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:57:13 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:57:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aviation incident landing site +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Flight (window number) +National Aeronautics and Space Administration (NASA) mission +crash site location +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the landing site that was the focus of a notable event or mission? +2025-04-03 at 08:57:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 08:57:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:57:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: recovery command module orbit or Earth-landing +2025-04-03 at 08:57:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 08:57:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the location of the Apollo spacecraft landing site? +2025-04-03 at 08:57:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 08:57:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:57:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Iwo Jima command module recovery +2025-04-03 at 08:57:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 08:57:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the approximate latitude and longitude of the Apollo 13 lunar module's separation point and impact location? +2025-04-03 at 08:57:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:57:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:57:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB impact location latitude and longitude, April 14, 1970 +2025-04-03 at 08:57:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 08:57:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:57:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:57:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, False, True, True] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:62 - Student lengths: [528, 579, 987, 546, 342, 255] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:64 - Average student length: 539.50 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 17.00 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_correctness:66 - Length ratio: 31.74 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.300 ± 0.328 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 1.60 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 0, 0, 0, 1, 4] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nTABLE 1O.3-I.- RECOVERY SUPPORT\n\nLanding area Supporta Remarks Number Unit La...', 'Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...', 'Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...'] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 08:57:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nTABLE 1O.3-I.- RECOVERY SUPPORT\n\nLanding area Supporta Remarks Number Unit La...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nIn prior lunar missions, the third stage has been separated from the spacecra...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...'] +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:57:27 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:57:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:57:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage signal intensity during ascent +2025-04-03 at 08:57:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 08:57:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 ascent stage impact signal shape and intensity" +2025-04-03 at 08:57:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 08:57:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:57:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: peak signal intensity Apollo 12 ascent stage vs Apollo 13 S-IVB impact +2025-04-03 at 08:57:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 08:57:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:36 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:57:36 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:57:36 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, False, True, True] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:62 - Student lengths: [376, 276, 836, 741, 449, 884] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:64 - Average student length: 593.67 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 14.00 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_correctness:66 - Length ratio: 42.40 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.125 ± 0.191 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 0.94 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 0, 0, 2, 0] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 08:57:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 08:57:36 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:57:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:57:40 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:57:40 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:57:40 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, True, True] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:62 - Student lengths: [401, 565, 285, 481, 565, 528] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:64 - Average student length: 470.83 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_correctness:66 - Length ratio: 78.47 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.000 ± 0.000 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.00 ± 0.00 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 08:57:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:57:40 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:57:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the primary objectives in a general context? +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: four primary objectives detailed objectives +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "project management frameworks with primary objectives explanation" or "project management framework with four primary objectives" +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what does primary objectives mean in the context of project management and unrelated fields? +2025-04-03 at 08:57:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 08:57:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:57:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "primary objectives crew activities flight mission control" +2025-04-03 at 08:57:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 08:57:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how many detailed objectives were derived from the four primary objectives listed as spacecraft and lunar surface objectives +2025-04-03 at 08:57:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 08:57:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:57:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:57:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:57:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:57:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:57:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:57:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:57:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:57:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:57:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:57:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:57:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:57:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:58:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:58:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no additional information needed for this answer +2025-04-03 at 08:58:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:58:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:58:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, True, False] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:62 - Student lengths: [761, 543, 188, 28, 795, 304] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:64 - Average student length: 436.50 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_correctness:66 - Length ratio: 54.56 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.329 ± 0.317 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.50 ± 3.50 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 1, 0, 3, 10] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: ++21.0 +3.0 +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:58:02 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:58:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"NASA Shuttle main engine 5 shutdown time after second stage jettison" + +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 SII staging timeline +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"pear-h Houston spacecraft mission engine 5 shutdown time after S-II staging +2025-04-03 at 08:58:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 08:58:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:58:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 crew entry join S-II push–off point +2025-04-03 at 08:58:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 08:58:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 32 minute mark Apollo 1 S-II staging +2025-04-03 at 08:58:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 08:58:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 engine 5 shutdown +2025-04-03 at 08:58:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:58:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 S-II engine 5 shutdown reference +2025-04-03 at 08:58:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:58:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apolo l3 crew timeline +2025-04-03 at 08:58:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 08:58:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module landing time +2025-04-03 at 08:58:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:58:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module jettison and landing +2025-04-03 at 08:58:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:58:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:58:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:58:21 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, True, True, True] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:62 - Student lengths: [489, 784, 1691, 346, 510, 528] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:64 - Average student length: 724.67 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_correctness:66 - Length ratio: 80.52 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.275 ± 0.340 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.87 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 8, 1, 0, 0] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 08:58:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nBecause of a sudden loss of pressure at approximately 56 hours from one of th...'] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 08:58:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 08:58:21 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:58:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:58:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar flyby mission purpose +2025-04-03 at 08:58:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 08:58:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar flyby mission purpose apollo 13 +2025-04-03 at 08:58:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:58:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar flyby mission purpose Apollo 13 training and science +2025-04-03 at 08:58:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:58:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar flyby mission purpose Apollo 13 testing and training +2025-04-03 at 08:58:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:58:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "unsuccessful lunar landing training mission" +2025-04-03 at 08:58:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 08:58:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lunar flyby mission geology experiment" +2025-04-03 at 08:58:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 08:58:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:58:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:58:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:58:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:62 - Student lengths: [581, 1954, 480, 285, 932, 858] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [58, 58, 58, 58, 58, 58] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:64 - Average student length: 848.33 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 58.00 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_correctness:66 - Length ratio: 14.63 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.158 ± 0.354 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.00 ± 2.24 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 0, 6] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.·��····· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 08:58:42 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 08:58:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA launch site USA electric field meters installed +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What electric field meters were installed in the area north and west of the Kennedy Space Center launch site? +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"electric field meter usage by NASA launch site" + +2025-04-03 at 08:58:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 08:58:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:58:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:58:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:58:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, False, False, True] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:62 - Student lengths: [470, 596, 623, 407, 544, 852] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:64 - Average student length: 582.00 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_correctness:66 - Length ratio: 145.50 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.183 ± 0.184 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.00 ± 1.41 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 4, 0, 1, 1] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 08:58:50 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 08:58:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 midcourse correction 3 ignition cutoff +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ignition cutoff third midcourse correction +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ignition cutoff third midcourse correction Apollo mission +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars Curiosity Rover third midcourse correction ignition cutoff Mission Profile 3 +2025-04-03 at 08:58:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-03 at 08:58:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ignition cutoff third midcourse correction apollo 11 +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third midcourse correction ignition cutoff +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 first midcourse correction ignition cutoff +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 third midcourse correction ignition time +2025-04-03 at 08:58:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:58:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 08:59:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third midcourse correction ignition cutoff apollo 11 +2025-04-03 at 08:59:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:59:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 first midcourse correction vs third midcourse correction ignition cutoff +2025-04-03 at 08:59:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:59:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:59:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apoll 11 third midcourse correction ignition time +2025-04-03 at 08:59:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 08:59:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third midcourse correction ignition cutoff tables IV +2025-04-03 at 08:59:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 08:59:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:59:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:59:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:59:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:62 - Student lengths: [241, 485, 490, 637, 452, 699] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:64 - Average student length: 500.67 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_correctness:66 - Length ratio: 100.13 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.404 ± 0.331 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 1.67 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [4, 2, 4, 3, 0, 0] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 08:59:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 08:59:06 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:59:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:59:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 6 after 40 seconds +2025-04-03 at 08:59:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:59:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical engineering site 6 negative field value after 40 seconds +2025-04-03 at 08:59:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:59:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:59:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electric field strength site 6 launch +2025-04-03 at 08:59:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 08:59:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: difference between launch exhaust and normal electric field magnitude at site 6 after 40 seconds +2025-04-03 at 08:59:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:59:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:59:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: open the information>block +2025-04-03 at 08:59:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:59:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cam fork disconnect electric field +2025-04-03 at 08:59:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 08:59:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: induced magnetic field valve opening +2025-04-03 at 08:59:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 08:59:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: remote compressor electric field interaction +2025-04-03 at 08:59:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:59:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:59:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:59:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:62 - Student lengths: [429, 360, 538, 423, 911, 351] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:64 - Average student length: 502.00 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_correctness:66 - Length ratio: 31.38 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.163 ± 0.230 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.00 ± 3.65 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [10, 0, 0, 0, 0, 2] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 08:59:25 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 08:59:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module power capacity lunar surface stay +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module life support system duration Apollo mission +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how many hours of support systems intended lunar module crew after landing +2025-04-03 at 08:59:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:59:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 08:59:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module power hours after lunar landing +2025-04-03 at 08:59:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module crew support duration after separation from command and service modules +2025-04-03 at 08:59:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:59:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 08:59:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power consumption energy hours lunar landing +2025-04-03 at 08:59:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power hours lunar operations +2025-04-03 at 08:59:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:59:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power duration extended mission operations +2025-04-03 at 08:59:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:59:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what hours did the lunar module support after the lunar landing +2025-04-03 at 08:59:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 08:59:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power hours after jettison +2025-04-03 at 08:59:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 08:59:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent energy ampere-hours nominals +2025-04-03 at 08:59:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:59:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 08:59:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 08:59:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1983, 379, 545, 495, 663, 774] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:64 - Average student length: 806.50 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_correctness:66 - Length ratio: 100.81 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.300 ± 0.350 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 2.85 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 0, 1, 2, 0, 0] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes , at wh...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes , at wh...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes , at wh...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\n1.0 SUMMARY\n\nThe Apollo l3 mission, planned as a lunar landing in the Fra Mau...'] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...'] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 08:59:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 08:59:46 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 08:59:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "antenna alignment and positioning distance from line of sight due to angle difference" +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna boresight axis line of sight difference between two sets of angles +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: comparing ground station antenna boresight angles and altitude differences +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: angle difference due to antenna boresight +2025-04-03 at 08:59:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 08:59:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Iwo Jima antenna distance from line of sight in degrees" +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 high-gain antenna alignment error calibration +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: distance due to antenna boresight angle difference in km using small angles approximation +2025-04-03 at 09:00:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 09:00:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:00:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 antenna alignment error 35 degrees off-line-of-sight +2025-04-03 at 09:00:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:00:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: distance from antenna boresight angle difference to line of sight in km, considering yaw and pitch components of 12° and 13°, respectively +2025-04-03 at 09:00:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:00:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:00:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 antenna misalignment precision required for 35 degrees off-line-of-sight +2025-04-03 at 09:00:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:00:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate approximate distance to line-of-sight using yaw and pitch angle differences, considering 18 km and 30 km sides, given omega and theta respectively +2025-04-03 at 09:00:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:00:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: line-of-sight distance to Earth using the average radius +2025-04-03 at 09:00:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:00:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: distance difference due to latitude, considering a line-of-sight angle and 4 minutes difference in latitude +2025-04-03 at 09:00:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:00:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate distance per degree and resulting distance from line of sight using pitch and yaw angle differences and 6371 km radius to Earth +2025-04-03 at 09:00:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:00:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:00:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:62 - Student lengths: [550, 557, 693, 894, 1170, 1394] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:64 - Average student length: 876.33 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_correctness:66 - Length ratio: 87.63 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.433 ± 0.362 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.33 ± 2.49 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 4, 1, 0, 0, 7] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:00:18 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:00:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot laboratory studies Apollo +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module testing laboratory results +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module Pilot laboratory studies +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot laboratory studies +2025-04-03 at 09:00:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot laboratory study results +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module postflight testing and training +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: post-flight testing primary Command Module Pilot +2025-04-03 at 09:00:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:00:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot postflight testing ASHUR's +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module performance testing and issues +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apex Cover Jettison System 0.05 ppt light +2025-04-03 at 09:00:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ + +2025-04-03 at 09:00:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot diagnostic test results +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module performance analysis March 1970 +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: entry monitor system guide +2025-04-03 at 09:00:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:00:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:00:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module performance evaluation April 1970 +2025-04-03 at 09:00:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo 14 mission report +2025-04-03 at 09:00:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:00:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module postflight evaluation April 1970 +2025-04-03 at 09:00:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll l0 mission laboratory studies +2025-04-03 at 09:00:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’�� 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:00:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Mission Plan +2025-04-03 at 09:00:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:00:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:00:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, True] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1461, 1201, 494, 484, 2038, 1221] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:64 - Average student length: 1149.83 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 22.00 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_correctness:66 - Length ratio: 52.27 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.504 ± 0.413 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.33 ± 2.69 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [4, 3, 6, 0, 7, 0] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:00:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe performance of the command and service module systems is discussed in thi...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe extensive testing and analyses and the consistency with which the postfli...'] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 09:00:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:00:47 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:00:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:00:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle astronaut physical examination +2025-04-03 at 09:00:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:00:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 postMission physical examination results +2025-04-03 at 09:00:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Mission Support Performance +2025-04-03 at 09:00:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:00:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 PostMission Physical Examination Results +2025-04-03 at 09:00:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:00:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Mission Support Performance +2025-04-03 at 09:00:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:00:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Astronaut Physical Condition Report +2025-04-03 at 09:00:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:00:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Astronaut Physical Examination Section Title +2025-04-03 at 09:00:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:00:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:00:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Medical Examination Supplement +2025-04-03 at 09:00:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 09:00:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Apollo 13 Medical Evaluation Report +2025-04-03 at 09:01:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:01:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:62 - Student lengths: [246, 380, 202, 294, 262, 442] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:64 - Average student length: 304.33 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 25.00 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_correctness:66 - Length ratio: 12.17 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.158 ± 0.354 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.50 ± 3.35 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 9, 0, 0] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...', 'Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...'] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +2025-04-03 at 09:01:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:01:02 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:01:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the orbital altitude of the pericynthion after the first major maneuver of a certain spacecraft? + + +2025-04-03 at 09:01:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:01:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Voyager 1 pericynthion altitude after major maneuver +2025-04-03 at 09:01:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:01:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:01:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Voyager 1 pericynthion altitude free-return trajectory maneuver time +2025-04-03 at 09:01:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:01:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Voyager 1 planned pericynthion orbital altitude free-return versus non-free-return trajectory +2025-04-03 at 09:01:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:01:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:13 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:13 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:13 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, True, False] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:62 - Student lengths: [637, 407, 375, 521, 411, 388] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:64 - Average student length: 456.50 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_correctness:66 - Length ratio: 57.06 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.167 ± 0.251 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 1.11 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 1, 0, 3, 0] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:01:13 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:01:13 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:01:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material is commonly used for the inner lining or backing of window shades? +2025-04-03 at 09:01:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 09:01:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of fabric is stitched to the inner surface of a window shade? +2025-04-03 at 09:01:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:01:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:62 - Student lengths: [354, 584, 201, 1116, 1236, 533] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:64 - Average student length: 670.67 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_correctness:66 - Length ratio: 55.89 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.083 ± 0.186 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.33 ± 0.75 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 0, 0, 0, 0, 0] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:01:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:01:22 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:01:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the throttle position for pressurization isolation solenoid on Boeing 737? +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "pressurization isolation solenoid throttle position control in automotive systems +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what throttle position controls pressurization isolation in an engine system +2025-04-03 at 09:01:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:01:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:01:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: at what throttle position is the pressurization isolation solenoid closed during descent on a 737? +2025-04-03 at 09:01:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when is the pressurization isolation solenoid closed at minimal throttle on Boeing 737 during descent? +2025-04-03 at 09:01:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: time in minutes before shutdown when throttle is 12% on Boeing 737 during which pressurization isolation solenoid is closed +2025-04-03 at 09:01:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Boeingshutdown engine time in minutes with suit pressure consistent with cabin pressure beforehand +2025-04-03 at 09:01:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:01:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle at 37% before Transearth Injection on Boeing 737 +2025-04-03 at 09:01:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:01:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, False, True, True] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:62 - Student lengths: [509, 232, 610, 916, 635, 427] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:64 - Average student length: 554.83 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 25.00 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_correctness:66 - Length ratio: 22.19 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.275 ± 0.340 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 2.13 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 0, 0, 1, 1, 0] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:01:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...'] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 09:01:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:01:38 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:01:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, False, False] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:62 - Student lengths: [509, 360, 1710, 437, 524, 542] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:64 - Average student length: 680.33 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_correctness:66 - Length ratio: 68.03 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.000 ± 0.000 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.00 ± 0.00 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 09:01:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:01:43 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:01:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "field-change and sferics detection lightning-like discharge launch site 5" +2025-04-03 at 09:01:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 09:01:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "calibrated electric field meters lightning detection lightning discharge site 5 launch report" +2025-04-03 at 09:01:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 09:01:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "sferics and field-change detector detection lightning site 5 launch specifics" +2025-04-03 at 09:01:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 09:01:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:01:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:01:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:01:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, False, False] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:62 - Student lengths: [536, 529, 346, 485, 626, 400] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:64 - Average student length: 487.00 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 2.00 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_correctness:66 - Length ratio: 243.50 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.108 ± 0.242 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 1.12 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 3, 0, 0, 0, 0] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:01:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:01:51 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:01:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 command and service module operations history +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module CSM 109 operations history Gemini 64 +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: refine search query +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module 109 history operations hawthorne california +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX CSM 109 operations history location +2025-04-03 at 09:01:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:01:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: CSM-109 evaluation Apollo 13 +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command service module history testing facility Downey California +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center operations history CSM-109 +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manned Spacecraft Center Houston Texas CSM-109 operations history +2025-04-03 at 09:02:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:02:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:02:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo CSM-109 manufacturer performance testing +2025-04-03 at 09:02:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:02:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Crew-4 operations manual +2025-04-03 at 09:02:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:02:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:02:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Downey California contractor facility Postflight CSM-109 testing +2025-04-03 at 09:02:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:02:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module postflight testing ASHUR 109007 109008 +2025-04-03 at 09:02:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:02:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: CSM-109 post-flight testing and inspection Downey California +2025-04-03 at 09:02:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:02:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:02:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:02:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1770, 591, 507, 632, 386, 585] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:64 - Average student length: 745.17 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 18.00 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_correctness:66 - Length ratio: 41.40 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.450 ± 0.280 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.50 ± 1.98 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 2, 4, 0, 1, 2] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nwas added for cleaning the suits before ingress, the bristles on the vacuum b...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nThe performance of the command and service module systems is discussed in thi...'] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:02:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:02:18 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:02:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the percentage of block material crushed when an O-ring enters a chamfer in breech assembly gun manufacturing +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the percentage of a breech block or receiver that is crushed by the O-ring as it enters the chamfer in a typical firearms breech assembly, assuming AVPS (Avtomatika Velloza Pažvaldinis Šatas) style breech assembly design? +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the percentage of crushed circular cross-section when O-ring enters chamfer +2025-04-03 at 09:02:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:02:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:02:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:02:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:02:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, False, True] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:62 - Student lengths: [338, 595, 217, 493, 659, 710] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:64 - Average student length: 502.00 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 2.00 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_correctness:66 - Length ratio: 251.00 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.175 ± 0.175 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 0.50 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 1, 1, 1, 0] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 09:02:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:02:27 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:02:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of pressure relief valves that require redesign +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve redesign automotive industrial applications +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "refrigerant valve redesign" +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What types of valves are prone to redesign due to energy efficiency or regulatory compliance in industries like oil, gas, and nuclear? +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of valve is commonly redesigned in modern technology? +2025-04-03 at 09:02:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postlanding vent valve redesign aerospace applications +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: The role of added compression in O-ring failure and recommended O-ring sizes for valves with high pressure and temperature applications +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascent stage tank shutoff valve design change +2025-04-03 at 09:02:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:02:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lock mechanism failure postlanding vent valve +2025-04-03 at 09:02:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascent stage tank shutoff valve O-ring redesign +2025-04-03 at 09:02:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:02:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supply valve oxygen isolation redesign +2025-04-03 at 09:02:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shutoff valve O-ring material replacement history +2025-04-03 at 09:02:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shutoff valve locking mechanism adjustment +2025-04-03 at 09:02:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 09:02:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shutoff valve locking mechanism operation +2025-04-03 at 09:02:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 09:02:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA-S-70-5841 shutoff valve locking mechanism +2025-04-03 at 09:02:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:02:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:02:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: two-cable valve locking mechanism interaction +2025-04-03 at 09:02:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:02:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:02:48 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:62 - Student lengths: [550, 1777, 188, 249, 724, 480] +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:64 - Average student length: 661.33 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_correctness:66 - Length ratio: 55.11 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.467 ± 0.292 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 2.98 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 9, 0, 1, 2, 3] +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 5.0 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.833 +2025-04-03 at 09:02:48 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:02:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Tank line burst cause in transport or storage context +2025-04-03 at 09:02:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 09:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cause of tank line burst +2025-04-03 at 09:02:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:02:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1967 accident investigation electrical conduit failure +2025-04-03 at 09:03:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:03:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (causes of tank line failure during M1 Abrams Main Battle Tank firing) +2025-04-03 at 09:03:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 postflight command module reaction control system testing +2025-04-03 at 09:03:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams tank electrical system failure causes) +2025-04-03 at 09:03:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 and Apollo 12 1969 testing results reaction control system +2025-04-03 at 09:03:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams tank tank line failure sequence of events) +2025-04-03 at 09:03:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent and Descent Propulsion System testing Apollo 12 discrepancy +2025-04-03 at 09:03:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams tank bay 4 insulation burning cause) +2025-04-03 at 09:03:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:03:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reliable sources Apollo 12 flight conditions +2025-04-03 at 09:03:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:03:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams tank gusset-4 breech assembly gas leak) +2025-04-03 at 09:03:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:03:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Thermal protection Apollo 12 lunar module operation +2025-04-03 at 09:03:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:03:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams tank gusset-4 breech assembly design improvements) +2025-04-03 at 09:03:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F���The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:03:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 batteries temperature anomalies +2025-04-03 at 09:03:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:03:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (M1 Abrams RCIV miswiring) +2025-04-03 at 09:03:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:03:23 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:03:23 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:03:23 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, True, True, False] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:62 - Student lengths: [281, 1757, 453, 778, 311, 1456] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [30, 30, 30, 30, 30, 30] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:64 - Average student length: 839.33 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 30.00 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_correctness:66 - Length ratio: 27.98 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.448 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 3.77 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 8, 0, 0, 0, 8] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:03:23 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:03:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:03:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of object 8591 in frame 13-60 in NASA's LIDAR datasets +2025-04-03 at 09:03:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraft， + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:03:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:03:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Latitude for object in frame 13-60-8591 +2025-04-03 at 09:03:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:03:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:03:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission landing position frame 13-60-8591 +2025-04-03 at 09:03:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:03:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:03:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing position +2025-04-03 at 09:03:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:03:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:03:35 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:03:35 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:03:35 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:62 - Student lengths: [648, 73, 334, 1018, 372, 238] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:64 - Average student length: 447.17 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 7.00 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_correctness:66 - Length ratio: 63.88 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.133 ± 0.298 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 1.49 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 4, 0] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraft， + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 09:03:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:03:35 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:03:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:03:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Primary Upper Atmosphere Flight Experience (PAUSE) crew health evaluations and NASA Long-Duration Spaceflight Study (NLOS) crew physical exams +2025-04-03 at 09:03:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:03:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:03:41 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:03:41 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:03:41 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, False, True, False] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:62 - Student lengths: [497, 463, 218, 565, 273, 571] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:64 - Average student length: 431.17 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 1.00 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_correctness:66 - Length ratio: 431.17 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 0, 0, 0] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...'] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 09:03:41 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:03:41 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:03:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:03:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 device purpose +2025-04-03 at 09:03:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +according to ground instructions. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:03:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 Area 51 purpose +2025-04-03 at 09:03:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:03:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: recovery operations Apollo 8 device Site 5 +2025-04-03 at 09:03:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:03:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Area 51 device site 5 operation +2025-04-03 at 09:03:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:03:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reaction Control System Apollo 8 +2025-04-03 at 09:03:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:03:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module separation system +2025-04-03 at 09:03:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:03:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reaction Control System Apollo 8 Site 5 +2025-04-03 at 09:03:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system service module separation procedure +2025-04-03 at 09:03:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:03:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Reaction Control System Ascent Propulsion System Apollo 8 Site 5 +2025-04-03 at 09:03:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module reaction control system site 5 +2025-04-03 at 09:03:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Reaction Control System Ascent Propulsion System Apollo 8 Site 5 +2025-04-03 at 09:03:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:03:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Reaction Control System component or system site 5 +2025-04-03 at 09:03:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:03:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service Module Reaction Control System Ascent Propulsion System Apollo 8 +2025-04-03 at 09:03:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:03:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:04:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:04:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:04:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, True] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:62 - Student lengths: [261, 460, 352, 746, 408, 535] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [46, 46, 46, 46, 46, 46] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:64 - Average student length: 460.33 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 46.00 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_correctness:66 - Length ratio: 10.01 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.448 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 3.08 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 7, 0, 6, 0, 0] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +according to ground instructions. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\naccording to ground instructions.\n------\nResult 2:\nThe Mission Control Center...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...'] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nSupport for the primary recovery area consisted of the prime recovery ship, U...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...'] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 09:04:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:04:01 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:04:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is X-Scale error units ppm +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What does the term "X-Scale factor error" refer to in the context of measuring signals or in a precision measurement application? +2025-04-03 at 09:04:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:04:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: curvature NOx pump scales ppm +2025-04-03 at 09:04:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:04:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor error in ppm definition +2025-04-03 at 09:04:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:04:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:04:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-scan error ppmlt +2025-04-03 at 09:04:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ +Result 2: +System testing with a similar antenna and electronics box showed RF signatures comparable to those observed in flight. This consistency was accomplished by placing the target inside the scan limits and the manual setting outside the scan limits. These two positions were separated approximately 35 degrees, which matched the actual angular separation experienced. Under these conditions, the antenna cycled between the target and the manual setting while operating in the auto-reacquisition mode and produced the cyclic RF signature. Since the inflight loss of signal to earth was not near the scan limit, the failure mechanism would be a shift in the Scan-limit function line. + +Elements in the scan-limit and scan-limit-warning circuit were shorted and opened to determine the effect on the scan-limit shift.The results of this test shifted the scan-limit functions but did not produce the necessary change in the scan-limit slope. Consequently, a failure in the electronic box is ruled out. +------ + +2025-04-03 at 09:04:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Meaning of X-Scale factor error pp solid-state accelerometer +2025-04-03 at 09:04:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:04:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:04:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-axis shift ppm scale factor +2025-04-03 at 09:04:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:04:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor error solid-state accelerometer specification +2025-04-03 at 09:04:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:04:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:04:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor error flight testing accelerometer +2025-04-03 at 09:04:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:04:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:04:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:04:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, True, True] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:62 - Student lengths: [311, 1638, 331, 571, 1697, 292] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:64 - Average student length: 806.67 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 1.00 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_correctness:66 - Length ratio: 806.67 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.350 ± 0.394 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.05 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 4, 1, 5, 0] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ +Result 2: +System testing with a similar antenna and electronics box showed RF signatures comparable to those observed in flight. This consistency was accomplished by placing the target inside the scan limits and the manual setting outside the scan limits. These two positions were separated approximately 35 degrees, which matched the actual angular separation experienced. Under these conditions, the antenna cycled between the target and the manual setting while operating in the auto-reacquisition mode and produced the cyclic RF signature. Since the inflight loss of signal to earth was not near the scan limit, the failure mechanism would be a shift in the Scan-limit function line. + +Elements in the scan-limit and scan-limit-warning circuit were shorted and opened to determine the effect on the scan-limit shift.The results of this test shifted the scan-limit functions but did not produce the necessary change in the scan-limit slope. Consequently, a failure in the electronic box is ruled out. +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:04:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:04:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:04:22 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:04:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Command Module Pilot two days before launch December 2022" +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot replacement astronaut Apollo moon mission +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: who was substituted as Command Module Pilot two days before launch in the Apollo space mission +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: who was replaced as Command Module Pilot two days before launch of a specific mission like Apollo 14 or another mission +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot substituted two days before Gemini launch +2025-04-03 at 09:04:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ + +2025-04-03 at 09:04:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Command Module Pilot Illness and Pre-Launch Replacement and Lunar Module Powering During Entry +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 transfer astronaut +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: who was substituted as Command Module Pilot of Apollo 13 two days before launch +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did NASA Gemini have crew substitution on launch day +2025-04-03 at 09:04:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:04:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Command Module Pilot name and reason for replacement +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module pilot replacement +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 original Command Module Pilot +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Who was substituted as Gemini Command Module Pilot before Gemini 1 launch due to rubella exposure +2025-04-03 at 09:04:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:04:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Charles "Pete" Conrad Apollo 10 Commander +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module pilot replacement +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission original crew +2025-04-03 at 09:04:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Charles "Pete" Conrad Apollo 10 +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module commander replacement +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: original Apollo 13 crew members +2025-04-03 at 09:04:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 crew medical replacement +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module commander +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew members besides James A. Lovell and John "Jack" Swigert +2025-04-03 at 09:04:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew medical evaluation Command Module Pilot substituted +2025-04-03 at 09:04:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Several days prior to laumch, the backup Lunar Module Pilot became sick with measles. Examinations of the prime crew indicated that the Command Module Pilot was not immune to the disease; therefore, the backup Command Module Pilot was substituted. The last 2 days prior to flight were devoted to integrated training among the three crew members, including the new Command Module Pilot. Flight results indicate that the last minute change of Command Module Pilots was practical and presented no training deficiencies, including readiness for the abort condition that occurred. + +8.2 PRELAUNCH PREPARATION +------ + +2025-04-03 at 09:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew +2025-04-03 at 09:04:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:04:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:04:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:04:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, False, True] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:62 - Student lengths: [721, 657, 368, 452, 794, 143] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:64 - Average student length: 522.50 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 20.00 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_correctness:66 - Length ratio: 26.12 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.562 ± 0.336 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.17 ± 3.02 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 8, 6, 1, 7, 3] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +Several days prior to laumch, the backup Lunar Module Pilot became sick with measles. Examinations of the prime crew indicated that the Command Module Pilot was not immune to the disease; therefore, the backup Command Module Pilot was substituted. The last 2 days prior to flight were devoted to integrated training among the three crew members, including the new Command Module Pilot. Flight results indicate that the last minute change of Command Module Pilots was practical and presented no training deficiencies, including readiness for the abort condition that occurred. + +8.2 PRELAUNCH PREPARATION +------ + +2025-04-03 at 09:04:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nEight days before flight, the primary Command Module Pilot was exposed to rub...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nLanding decelerations were mild in comparison to Apollo 8, and the spacecraft...', 'Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...'] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:04:47 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:04:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "computer-displayed inertial velocity at cutoff" or "computer-displayed inertial velocity at cutoff value" flight or spacecraft +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: computer display inertial velocity at cutoff event +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: computer-displayed inertial velocity UFO incident G Palestine +2025-04-03 at 09:04:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:04:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff inertial velocity +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ap17 mission inertial velocity at cutoff +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: second midcourse correction ignition cutoff +2025-04-03 at 09:04:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:04:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: second midcourse correction data velocity +2025-04-03 at 09:04:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:04:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:04:59 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:04:59 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:04:59 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, False, True, False] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:62 - Student lengths: [656, 260, 211, 513, 499, 384] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:64 - Average student length: 420.50 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_correctness:66 - Length ratio: 32.35 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.275 ± 0.280 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.17 ± 1.21 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 2, 0, 0, 3] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...'] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...'] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:04:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\ninjection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866...', 'Result 1:\ndata. Following this maneuver, a series of earth photographs were taken for l...'] +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:04:59 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:05:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 202 SC-011 accident August 25 1966" +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission failure cause August 25 1966 +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: as 202 sc 011 incident august 25 1966 type of entry +2025-04-03 at 09:05:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:05:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 7 launch August 25 1966" +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 mission entry причина +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo AS-202 SC-011 re-entry type +2025-04-03 at 09:05:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:05:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 1966 launch accident August" +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 mission August 25 entry details +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo AS-202 SC-011 spacecraft re-entry type 1966 +2025-04-03 at 09:05:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 1 fire August 1966" +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 accident August 25 details +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 201 Saturn C-1 re-entry characteristics August 25 1966 +2025-04-03 at 09:05:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:05:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact details +2025-04-03 at 09:05:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:05:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:05:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:05:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:05:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 0/6 answers correct +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.00 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:62 - Student lengths: [447, 265, 314, 295, 1099, 747] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [39, 39, 39, 39, 39, 39] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:64 - Average student length: 527.83 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 39.00 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_correctness:66 - Length ratio: 13.53 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.425 ± 0.428 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 2.19 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 4, 5, 0, 4] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:05:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...'] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:05:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...'] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:05:18 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:05:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "effects of escaped propellant gas at launch or aerospace equipment +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "chemical contamination from propellant leakage on adjacent equipment" +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " rocket propellant gas escape damage assessment on adjacent equipment or components" or "investigation of hydrogen propellant gas leak on adjacent equipment" +2025-04-03 at 09:05:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:05:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "breech-plenum interface failure consequences on adjacent equipment in aerospace or spacecraft systems +2025-04-03 at 09:05:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "electrical component contamination from propellant gas failure" +2025-04-03 at 09:05:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:05:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:05:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polymide vs metal erosion from high-temperature propellant gas leak +2025-04-03 at 09:05:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:05:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:05:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: consequences of gas leakage at breech-plenum interface +2025-04-03 at 09:05:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:05:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fire-related residues or combustion byproducts +2025-04-03 at 09:05:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:05:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:05:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal decomposition products from overheating electrical conduits +2025-04-03 at 09:05:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 09:05:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:05:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:05:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:05:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:62 - Student lengths: [660, 290, 1016, 370, 984, 1964] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:64 - Average student length: 880.67 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_correctness:66 - Length ratio: 220.17 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.283 ± 0.350 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.13 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 6, 0, 2, 2, 0] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:05:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:05:37 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:05:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Terra Nova-4 environmental control system duration +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"how long did the environmental control system provide a habitable environment for the Apollo 8 mission" + +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars mission with a crew that lived in a controlled environment for several hours or days +2025-04-03 at 09:05:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ + +2025-04-03 at 09:05:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:05:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:05:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, False] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:62 - Student lengths: [454, 335, 308, 905, 478, 511] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:64 - Average student length: 498.50 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 2.00 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_correctness:66 - Length ratio: 249.25 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.171 ± 0.171 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.83 ± 1.07 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 3, 0, 0, 0] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ + +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 09:05:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:05:47 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:05:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what led to the shift from automatic to manual systems in industry? +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what triggered the switch to manual system in aviation +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: trigger manual system +2025-04-03 at 09:05:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-03 at 09:05:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Space Shuttle crew manual override after oxygen tank anomaly +2025-04-03 at 09:05:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank anomaly switch to manual system +2025-04-03 at 09:05:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:05:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:05:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module manual override oxygen tank malfunction +2025-04-03 at 09:05:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 09:05:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 manual override decision oxygen tank malfunction +2025-04-03 at 09:06:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:06:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission manual override life support allocation +2025-04-03 at 09:06:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:06:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 manual override power conservation and battery usage +2025-04-03 at 09:06:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module battery operation energy management +2025-04-03 at 09:06:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 manual override energy management false battery 2 malfunction +2025-04-03 at 09:06:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module battery malfunction led to manual override +2025-04-03 at 09:06:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:06:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:06:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:62 - Student lengths: [945, 578, 988, 1941, 970, 505] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:64 - Average student length: 987.83 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 28.00 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_correctness:66 - Length ratio: 35.28 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.300 ± 0.350 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.00 ± 3.21 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 0, 0, 9, 1, 0] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nAt the time of the oxygen tank incident, three events took place that affecte...'] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe two tanks cortaining cryogenic oxygen, used for _fuel cell operation and ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe designs of other Apollo batteries have been reevaluated, and all are cons...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...'] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 09:06:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:06:09 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:06:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:06:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much ampere hours were remaining in the lunar module batteries at undocking? Apollo 11 +2025-04-03 at 09:06:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Apollo lunar module power status during undocking +2025-04-03 at 09:06:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:06:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "power reserve lunar module undocking Apollo mission +2025-04-03 at 09:06:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:06:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Apollo 11 lunar module descent battery state during undocking +2025-04-03 at 09:06:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Lunar Module Descent Battery State During Undocking Apollo 11 NASA Report +2025-04-03 at 09:06:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:06:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:06:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:62 - Student lengths: [437, 705, 623, 540, 732, 464] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:64 - Average student length: 583.50 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 3.00 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_correctness:66 - Length ratio: 194.50 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.192 ± 0.301 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.83 ± 1.46 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 1, 4] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:06:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:06:22 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:06:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"meru robot technical documentation" + +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: p Bendulum acceleration Z-direction unit +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration drift in mERU/g Z-direction +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration drift z-axis mERU/g +2025-04-03 at 09:06:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 09:06:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"meru acceleration error specifications" + +2025-04-03 at 09:06:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gaussian vector, inertial measurements lunar module +2025-04-03 at 09:06:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:06:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:06:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Z-axis accelerometer bias) saturation mERG +2025-04-03 at 09:06:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:06:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:06:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:06:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:06:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:62 - Student lengths: [167, 76, 945, 423, 374, 421] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:64 - Average student length: 401.00 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_correctness:66 - Length ratio: 100.25 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.304 ± 0.237 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 1.80 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 0, 3, 0, 1, 1] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...'] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nTime hr:min Optian code Star used Ster angle aifference, deg Gyro torquing an...'] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 09:06:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:06:39 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:06:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: _weather conditions before Apollo 13 launch +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How was the weather forecast for the Apollos 11 (since it has errors if I omit the 11 and try using 13 instead) launch described? + + +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather April 11 1970 +2025-04-03 at 09:06:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:06:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: conditions during the cold front on April 11, 1970 +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the forecasted weather situation on April 11, 1970, for the launch of Apollo 13, specifically the time around liftoff and its relation to pre-launch issues at Site 5? + + +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather April 11 1970 +2025-04-03 at 09:06:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:06:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the weather conditions at Cape Canaveral on April 11, 1970, around the time of the Apollo 13 launch? + + +2025-04-03 at 09:06:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather forecast April 11 1970 +2025-04-03 at 09:06:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:06:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:06:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What weather warnings or forecasts were available for Cape Canaveral on April 11, 1970, as part of the Apollo 13 mission's launch preparations? + + +2025-04-03 at 09:06:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:06:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather report April 11 1970 +2025-04-03 at 09:06:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:06:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:06:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were there any storm warnings issued for the Cape Canaveral area around April 10-11, 1970? + + +2025-04-03 at 09:06:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:06:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather report April 11 1970 +2025-04-03 at 09:06:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:06:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:06:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were there any significant weather events or forecasts related to the Atlantic coastal area around April 1970 that could impact a launch from Cape Canaveral? + + +2025-04-03 at 09:06:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:06:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission launch day weather reports +2025-04-03 at 09:06:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:06:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the final weather forecast or operating conditions report for Cape Canaveral AFB, Florida, submitted on April 11, 1970? + + +2025-04-03 at 09:07:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:07:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: April 12 cold front and Apollo 13 launch +2025-04-03 at 09:07:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:07:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11-12 cold front weather April 11 1970 +2025-04-03 at 09:07:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:07:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:07:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:07:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, False, True] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:62 - Student lengths: [974, 610, 327, 892, 855, 671] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:64 - Average student length: 721.50 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_correctness:66 - Length ratio: 90.19 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.400 ± 0.427 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.83 ± 3.39 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 0, 7, 0, 8] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:07:06 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:07:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:07:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "initial launch azimuth of a typical spacecraft" +2025-04-03 at 09:07:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:07:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial azimuth launch vehicle map +2025-04-03 at 09:07:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:07:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 launchpad bearing or azimuth at liftoff" +2025-04-03 at 09:07:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:07:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:16 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:07:16 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:07:16 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, False, False] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:62 - Student lengths: [420, 578, 569, 729, 604, 592] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:64 - Average student length: 582.00 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 24.00 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_correctness:66 - Length ratio: 24.25 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.142 ± 0.205 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 0.76 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 1, 0, 0, 0, 0] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:07:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:07:16 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:07:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:07:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hard uso annual-tubing- timothy severins joshua pritchard balloon flight +2025-04-03 at 09:07:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:07:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the launch times of the first two balloons launch in history? +2025-04-03 at 09:07:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:07:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 equipment launch times April 11 1970 +2025-04-03 at 09:07:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 launch timing April 11 1970 +2025-04-03 at 09:07:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 CSM liftoff time April 11 1970 +2025-04-03 at 09:07:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo CSM liftoff confirmation April 11 1970 +2025-04-03 at 09:07:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:30 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:07:30 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:07:30 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.17 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:62 - Student lengths: [517, 28, 116, 227, 245, 102] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:64 - Average student length: 205.83 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_correctness:66 - Length ratio: 15.83 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.200 ± 0.348 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.17 ± 1.86 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 5, 0, 0, 0] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:07:30 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 09:07:30 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:07:30 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:07:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Y-axis error magnitude discrepancy analysis context +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "discrepancy in x-axis error magnitude" +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the Y-axis error magnitude issue in metrology? +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Y-axis error magnitude discrepancy instrumentation calibration physics engineering +2025-04-03 at 09:07:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 09:07:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:07:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "x-axis error sources" +2025-04-03 at 09:07:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 platform coarse alignment error counter initialization +2025-04-03 at 09:07:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:07:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 null bias drift timing and platform alignment +2025-04-03 at 09:07:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:07:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 null bias drift calculation platform alignment +2025-04-03 at 09:07:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:07:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:07:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:07:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:62 - Student lengths: [367, 549, 586, 588, 665, 443] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [36, 36, 36, 36, 36, 36] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:64 - Average student length: 533.00 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 36.00 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_correctness:66 - Length ratio: 14.81 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.279 ± 0.205 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.05 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 2, 0, 0, 1, 6] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...'] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:07:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...'] +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:07:43 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:07:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the earliest recorded data point in human history +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: value="first recorded data point time range" search term +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the date of the earliest recorded astronomical observation or systematic timekeeping? +2025-04-03 at 09:07:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 09:07:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:07:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 impact time range +2025-04-03 at 09:07:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:07:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Time range of first seismic signal recorded by seismic instrument during Apollo 12 mission +2025-04-03 at 09:07:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:07:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 SIVB impact time +2025-04-03 at 09:07:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:07:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of the first seismic signal recorded by the Apollo 12 seismometer +2025-04-03 at 09:07:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:07:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of the seismic signal after initial peak to 4-hour recording period during Apollo 12 mission +2025-04-03 at 09:07:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:07:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of seismic signal after 10 minutes of initial peak signal +2025-04-03 at 09:07:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 09:07:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of initial stage of S-IVB impact seismic signal +2025-04-03 at 09:07:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:07:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:07:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Characteristics of initial stage of S-IVB impact seismic signal +2025-04-03 at 09:07:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:07:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Velocity of initial stage of S-IVB impact seismic signal +2025-04-03 at 09:08:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:08:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of initial stage of S-IVB impact seismic signal +2025-04-03 at 09:08:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Duration of signal from beginning to peak value for S-IVB impact +2025-04-03 at 09:08:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:08:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:08:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:08:03 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, False, True] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:62 - Student lengths: [262, 28, 577, 1809, 274, 915] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:64 - Average student length: 644.17 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_correctness:66 - Length ratio: 128.83 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.325 ± 0.368 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.33 ± 3.59 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 1, 0, 10, 0, 0] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:08:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...'] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:08:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nFollowing translunar injection, earth weather photography was conducted for a...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...'] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:08:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:08:03 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:08:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why did the failure of the service propulsion auxiliary propellant gaging system impact the performance of a space mission? +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "service propulsion auxiliary propellant gaging system mission performance" +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft propulsion system failure impact assessment" +2025-04-03 at 09:08:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:08:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the failure of the service propulsion auxiliary propellant gaging system cause a loss of primary oxygen in the service module? +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "auxiliary propellant gaging system failure impact on mission parameters" +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle auxiliary propellant gaging system failure impact assessment +2025-04-03 at 09:08:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the failure of the service propulsion auxiliary propellant gaging system cause any delays or changes in the mission plan? +2025-04-03 at 09:08:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "auxiliary propellant gaging system failure propellant usage discrepancy" +2025-04-03 at 09:08:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: was the failure of the service propulsion auxiliary propellant gaging system a contributing factor to the abort of the mission due to primary oxygen loss? +2025-04-03 at 09:08:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "propellant usage increase due to propulsive venting" +2025-04-03 at 09:08:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 09:08:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:08:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "high flow rate reaction control system propellant usage" +2025-04-03 at 09:08:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:08:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "reaction control system propellant usage specific event propulsive venting" +2025-04-03 at 09:08:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:08:24 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:08:24 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:08:24 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, True, False] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:62 - Student lengths: [517, 684, 633, 527, 636, 429] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:64 - Average student length: 571.00 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 2.00 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_correctness:66 - Length ratio: 285.50 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.375 ± 0.398 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.00 ± 2.31 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 4, 6, 0, 2, 0] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 09:08:24 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:08:24 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:08:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection pericynthion altitude +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Translunar injection pericynthion altitude Apollo mission" +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection altitude +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection pericynthion space mission +2025-04-03 at 09:08:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:08:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: parsec height achieved during translunar injection for apollo 11 +2025-04-03 at 09:08:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection altitude +2025-04-03 at 09:08:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:08:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:08:34 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:08:34 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:08:34 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:62 - Student lengths: [361, 520, 226, 491, 169, 598] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:64 - Average student length: 394.17 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 11.00 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_correctness:66 - Length ratio: 35.83 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.267 ± 0.207 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.17 ± 0.90 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 0, 1, 2, 2, 0] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:08:34 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 09:08:34 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:08:34 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:08:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: The dimensions of the Starship's Command and Service Module (CSM) are available through official SpaceX resources, but I will look up the typical dimensions for an object of that composition to make an educated estimate. +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft Command Service Module moment inertia +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service module iss mass and dimensions +2025-04-03 at 09:08:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:08:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll look for information on the moment of inertia of a CSM similar to the one described in the manual, but also try to find information specific to the Starship's Command and Service Module. +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mass Command Service Module +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service module iss dimensions +2025-04-03 at 09:08:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:08:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll also try to find information on the spacecraft's attitude control system to determine the axis of rotation. +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mass of a spacecraft's Command Module +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service module iss dimensions +2025-04-03 at 09:08:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:08:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll search for technical documents or publications that discuss the Starship's Command and Service Module, specifically its mass, dimensions, and any relevant technical specifications that might be necessary to calculate its moment of inertia. +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia Command and Service Module +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service module iss length and width +2025-04-03 at 09:08:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:08:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll try a different approach and search for mission design requirements documents or technical manuals for the Starship that might include information on the Command and Service Module's design, including its size, mass, and any relevant technical specifications. +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of a similar spacecraft module +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service module iss dimensions technical specifications +2025-04-03 at 09:08:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:08:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll search for the full report MSC-02680 to see if there's any relevant information that I missed, specifically section A.5 Mass Properties, which should contain information on the mass and possibly some related specifications that might help calculate the moment of inertia of the Command and Service Module. +2025-04-03 at 09:09:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia formula for a cylinder +2025-04-03 at 09:09:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:09:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:09:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll search for detailed information on how to calculate the moment of inertia of an object given its mass and physical dimensions, which might help me calculate the CSM's moment of inertia using the values provided in Table A-I. +2025-04-03 at 09:09:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 09:09:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of a cylinder formula +2025-04-03 at 09:09:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:09:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:09:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: I'll search for documents related to the Starship's technical specifications, specifically the Command and Service Module's design, to see if I can find any information on its mass, dimensions, and shape, which could be used to calculate its moment of inertia. +2025-04-03 at 09:09:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:09:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:09:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, True, False] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:62 - Student lengths: [2458, 733, 696, 2049, 854, 926] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:64 - Average student length: 1286.00 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_correctness:66 - Length ratio: 257.20 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.354 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.67 ± 4.82 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 11, 7, 10, 0] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:09:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nDESCRIPTIONS·········· A-1 A...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe service module first app...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe service module first app...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe service module first app...', 'Result 1:\nThe service module first appeared in the docking window at a distance of abou...'] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 09:09:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:09:08 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:09:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 mission manual antenna adjustments +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Command Module Pilot antenna angle adjustment Apollo 11 lunar module" +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Command Module Pilot antenna setting +2025-04-03 at 09:09:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:09:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:09:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Command Module Pilot antenna settings angle of pitch +2025-04-03 at 09:09:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:09:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Mission Command Module Pilot S-band communications antenna adjustments +2025-04-03 at 09:09:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:09:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:09:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Mission Manual Command Module Pilot Antenna Angle of Pitch Adjustment +2025-04-03 at 09:09:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:09:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 antenna beam switch manual adjustments +2025-04-03 at 09:09:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:09:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:09:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 pitching adjustments to antenna direction +2025-04-03 at 09:09:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:09:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:09:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:09:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:09:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, True, True, True] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:62 - Student lengths: [562, 242, 352, 607, 397, 747] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:64 - Average student length: 484.50 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_correctness:66 - Length ratio: 48.45 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.300 ± 0.328 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 1.60 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 0, 0, 1, 0, 4] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...'] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:09:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:09:22 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:09:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo missions emergency abort +2025-04-03 at 09:09:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission emergency abort +2025-04-03 at 09:09:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:09:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:09:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar orbit splashdown +2025-04-03 at 09:09:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:09:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:09:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:09:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:09:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:62 - Student lengths: [548, 805, 1220, 614, 785, 731] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:64 - Average student length: 783.83 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_correctness:66 - Length ratio: 87.09 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.142 ± 0.205 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 0.76 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 0, 0, 0, 2] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:09:32 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:09:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "iPhone 1 battery capacity +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much energy did the original Tesla Roadster have in its three entry batteries? +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the battery capacities of the original iPhone 12 +2025-04-03 at 09:09:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:09:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: iPhone 1 battery capacity original +2025-04-03 at 09:09:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Find the total energy in the Lunar Module's batteries at undocking. +2025-04-03 at 09:09:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:09:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:09:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command and Service Module battery capacity +2025-04-03 at 09:09:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:09:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:09:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:09:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:09:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, False, False, True] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1114, 135, 445, 678, 332, 378] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:64 - Average student length: 513.67 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 15.00 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_correctness:66 - Length ratio: 34.24 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.250 ± 0.265 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.00 ± 1.15 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 1, 2, 0, 0, 0] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 09:09:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:09:43 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:09:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo biomedical monitoring signals and command module capacity +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many vital signs northing monitored in the lunar module? +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical number of biomedical signals that can be monitored in a typical space suit or life support system on the Lunar Module? +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo biomedical signals monitored +2025-04-03 at 09:09:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:09:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module biomedical monitoring signal capabilities +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many simultaneous biomedical signals could the Lunar Module monitor? +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the number of physiological signals or channels that have simultaneous monitoring capabilities in a Lunar Module astronaut? +2025-04-03 at 09:09:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:09:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Apollo biomedical signal storage and power consumption +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Does the Lunar Module's onboard computer document a maximum number of simultaneous mandatory or recommended biomedical signals to monitor during operations? +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many channels or signals of a specific type (e.g., physiological signals like respiration or blood pressure) can be monitored simultaneously by the Lunar Module's life support system? +2025-04-03 at 09:09:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:09:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module power allocation for biomedical equipment +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can the Lunar Module monitor a limited number of signals while still allowing for crew safety? +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is there any information on the number of channels for a typical lunar module's life support system, and are there any sources that mention the monitoring of respiration or blood pressure? +2025-04-03 at 09:09:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:09:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo biomedical monitoring system power requirements +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can the Lunar Module manage fewer signals due to command module separation? +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can any space mission or space suit documents or reports discuss the dynamic allocation or prioriory of channels utilized by the Apollo Lunar Module for its life support system? +2025-04-03 at 09:10:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:10:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module undocking procedure focuses on whereas the command module is powered up +2025-04-03 at 09:10:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:10:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Are any mission documentation, technical manuals, or guidelines that reference the specific capabilities of the Apollo Lunar Module's life support system available? +2025-04-03 at 09:10:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:10:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:10:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the Lunar Module's single electrocardiogram signal the primary limiting factor? +2025-04-03 at 09:10:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-03 at 09:10:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can the Apollo 13 mission technical data or documentation be found in resources such as NASA's official Apollo mission archives? +2025-04-03 at 09:10:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:10:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:10:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can any NASA documents or technical reports, such as the Apollo Spacecraft Design Requirements or Manual, provide information on the life support system's capabilities, including the number of channels or signals that can be monitored? +2025-04-03 at 09:10:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:10:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:10:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:62 - Student lengths: [853, 513, 1777, 1321, 453, 860] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:64 - Average student length: 962.83 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 3.00 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_correctness:66 - Length ratio: 320.94 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.533 ± 0.433 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.50 ± 3.30 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 7, 8, 0, 0, 1] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:10:12 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:10:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: primary mission decision cryogenic oxygen tank problem +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: .primary mission decision after cryogenic oxygen tank failure tanks military +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic oxygen tank failure mission decision +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space mission cryogenic oxygen tank failure +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Doolittle Raid Mission Decision after cryogenic oxygen tank problem +2025-04-03 at 09:10:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:10:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:10:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 oxygen tank failure mission decision +2025-04-03 at 09:10:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 oxygen tank failure emergency abort decision +2025-04-03 at 09:10:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:24 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:10:24 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:10:24 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:62 - Student lengths: [912, 434, 652, 543, 489, 487] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [84, 84, 84, 84, 84, 84] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:64 - Average student length: 586.17 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 84.00 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_correctness:66 - Length ratio: 6.98 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.325 ± 0.191 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 0.94 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 3, 0, 1, 2] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:10:24 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nCryogenic storage system operation was satisfactory until 46:40:09, when the ...'] +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:10:24 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:10:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System function" +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System purpose and function" +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon lunar module control system purpose +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module reaction control system purpose +2025-04-03 at 09:10:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:10:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:10:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System role in Apollo mission management" +2025-04-03 at 09:10:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo Mission Data Reviews of Lunar Module and Service Module Reaction Control Systems" +2025-04-03 at 09:10:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission evaluation reports performance of reaction control system" +2025-04-03 at 09:10:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:10:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module reaction control system post-flight testing and evaluation" +2025-04-03 at 09:10:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System safety features and systems" +2025-04-03 at 09:10:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:10:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:10:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:10:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, True] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:62 - Student lengths: [625, 656, 726, 551, 721, 892] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:64 - Average student length: 695.17 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 28.00 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_correctness:66 - Length ratio: 24.83 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.333 ± 0.317 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.50 ± 2.06 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 6, 1, 0, 0, 1] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...'] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...'] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:10:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...'] +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:10:37 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:10:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +flight crew Pago Pago, Samoa destination +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where were the flight crew flown to after being flown from Pago Pago? air india 101, flight 103 +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where was the Pago Pago flight crew flown to after an incident +2025-04-03 at 09:10:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:10:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +crew from Iwo Jima to Pago Pago, Samoa destination +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where were Air India Flight 101 and 103 flight crew flown to after being diverted from Pago Pago, Samoa? +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Iwo Jima to Samoa flying crew transport +2025-04-03 at 09:10:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Pago Pago, Samoa crew flight destination +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where were the crew of Air India Flight 101 and 103 flown to after a diversion +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: US military transport of personnel from Iwo Jima to Pago Pago, Samoa +2025-04-03 at 09:10:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +flight crew from Pago Pago, Samoa, to +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Destination of Air India Flight 101 and 103 after Pago Pago, Samoa +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Iwo Jima recovery mission April 17, 1970, to Pago Pago, Samoa +2025-04-03 at 09:10:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:10:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Iwo Jima crew departure destination +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where were Air India Flight 101 and 103 crew flown after landing in Pago Pago, Samoa +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: C-l4l aircraft flight from Iwo Jima to Pago Pago, Samoa +2025-04-03 at 09:10:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:10:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Iwo Jima crew from Pago Pago, Samoa destination +2025-04-03 at 09:10:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:10:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Air India Flight 101 and 103 diversion destination +2025-04-03 at 09:10:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:10:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:11:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Iwo Jima departure from Pago Pago, Samoa +2025-04-03 at 09:11:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Air India Flight 101 and 103 attempted diversion +2025-04-03 at 09:11:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:11:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:11:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Pago Pago, Samoa to Hawaii +2025-04-03 at 09:11:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Air India Flight 101 and 103 recovery operations +2025-04-03 at 09:11:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:11:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where was the crew of Air India Flight 101 and 103 after a diversion +2025-04-03 at 09:11:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:11:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Air India Flight 101 and 103 diversion +2025-04-03 at 09:11:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:11:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:11:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, False] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:62 - Student lengths: [696, 718, 183, 987, 441, 387] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:64 - Average student length: 568.67 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_correctness:66 - Length ratio: 94.78 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.475 ± 0.475 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.83 ± 4.10 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 8, 10, 5, 0] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:11:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:11:06 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:11:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are used for interior components of spacecraft where extreme temperatures and radiation are experienced? +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are currently being researched or used for the interior of breech plenum areas in spacecraft design? +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft breech plenum area thermal insulation materials" +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space aircraft breech plenum lining material +2025-04-03 at 09:11:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials used in Apollo 11 spacecraft interior components +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are being researched or developed for high-temperature, high-pressure applications in space, and are compatible with the properties required for a failure of O-ring seals in such conditions? +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "tapioca-like materials for aerospace applications" +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: sheet material for aerospace high temperature and vacuum breech plenum +2025-04-03 at 09:11:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials used in Apollo 11 instrument packages or electronics equipment breech plenum area +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are suitable for high-temperature and high-pressure applications, and are capable of maintaining seals in conjunction with stainless steel? +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "teflon alternatives for aerospace materials in breech plenum areas" +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide sheet material properties for high temperature and pressure applications in aerospace +2025-04-03 at 09:11:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials used in Apollo 11 SIVB lunar impact experiment module breech assembly +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are used in conjunction with stainless steel for gas containment and sealing applications in cryogenic systems? +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "polyimide vs Teflon for thermal protection in aerospace applications" +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide sheet material properties for aerospace applications used in Apollo spacecraft +2025-04-03 at 09:11:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials used in SIVB lunar impact experiment and resistant to pressure and gas leaks +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials are used in conjunction with stainless steel for high-pressure gas containment and sealing applications in cryogenic systems, with a focus on maintaining hermetic seals? +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "polyimide thermal insulation performance in extreme temperatures" +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: material specifications polyimide used in Apollo spacecraft breech plenum +2025-04-03 at 09:11:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spray foam alternatives used for insulation in spacecraft +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Advanced materials used for high-pressure seals in cryogenic applications, compatible with stainless steel +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Material properties of Mylar insulation in high-temperature and shock-resistant applications" +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide material used in Apollo breach plenum +2025-04-03 at 09:11:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials similar to spray foam used in S-Il stage +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Mylar insulation properties in high-pressure and high-temperature applications" +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type of polyimide material used in Apollo spacecraft +2025-04-03 at 09:11:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type and specifications of polyimide used in Apollo space program +2025-04-03 at 09:11:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of polyimide film used in aerospace applications +2025-04-03 at 09:11:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:11:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:11:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, True, False, False] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:62 - Student lengths: [676, 1759, 954, 873, 342, 1776] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:64 - Average student length: 1063.33 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_correctness:66 - Length ratio: 118.15 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.554 ± 0.426 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.00 ± 3.65 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 8, 0, 6, 7, 9] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 09:11:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...'] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:11:39 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:11:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space mission task verification simulator +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"NASA simulation task verification for space mission" + +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing simulator preflight verification task check + + +2025-04-03 at 09:11:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:11:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:11:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 lunar landing procedures simulator verification + + +2025-04-03 at 09:11:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:11:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection maneuver difficulties Apollo 8 + + +2025-04-03 at 09:11:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:11:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection maneuver Apollo 8 abort + + +2025-04-03 at 09:11:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:11:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection maneuver abort handling +2025-04-03 at 09:11:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:11:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:11:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection passive thermal control mode challenges +2025-04-03 at 09:11:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:11:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection passive thermal control mode advantages +2025-04-03 at 09:12:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:12:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection engine shutdown benefits + + +2025-04-03 at 09:12:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:12:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:12:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:12:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:62 - Student lengths: [468, 306, 575, 321, 857, 2017] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:64 - Average student length: 757.33 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_correctness:66 - Length ratio: 75.73 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.196 ± 0.200 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 3.98 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 0, 1, 11] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:12:02 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:12:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:12:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: can pressure in vacuum jacket be checked after pumpdown stabilization phase +2025-04-03 at 09:12:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 09:12:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum jacket pressure measurement after pumpdown 2 to 3 weeks +2025-04-03 at 09:12:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 09:12:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:12:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:12:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, True, True] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:62 - Student lengths: [763, 1044, 1184, 985, 535, 746] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:64 - Average student length: 876.17 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_correctness:66 - Length ratio: 73.01 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.083 ± 0.186 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.33 ± 0.75 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 0, 0, 0, 0, 0] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 09:12:12 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:12:12 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:12:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: LOLA Space Thermal Control System antenna passive mode +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: satellite antenna passive thermal control temperature range +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range antenna passive thermal control mode +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range of an antenna in a specific high-temperature application +2025-04-03 at 09:12:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:12:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range of passive thermal control antenna +2025-04-03 at 09:12:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 09:12:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range for wide medium and narrow beam antenna in space mission +2025-04-03 at 09:12:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:12:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:12:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature range passive mode engineering reference +2025-04-03 at 09:12:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic storage temperatures related to antennas in spacecraft +2025-04-03 at 09:12:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:12:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:12:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range S-band steerable antenna passive mode engineering reference +2025-04-03 at 09:12:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:12:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range for high-gain antenna cold temperatures for communication equipment +2025-04-03 at 09:12:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:12:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 2-lo foot receiving antenna S-band communication +2025-04-03 at 09:12:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-gain antenna temperature range near -330°C for communication equipment +2025-04-03 at 09:12:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:12:36 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:12:36 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:12:36 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, False, True, True] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1247, 143, 898, 494, 204, 647] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [29, 29, 29, 29, 29, 29] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:64 - Average student length: 605.50 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 29.00 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_correctness:66 - Length ratio: 20.88 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.337 ± 0.324 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.00 ± 3.56 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 5, 0, 1, 10] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:12:36 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\n14.l.4 High-Gain Antenna Acquisition Problem\n\nPrior to the television transmi...', 'Result 1:\nWhen the transmission was switched from the omnidirectional antenna to the ma...', 'Result 1:\nConsumable quantities in the cryogenic storage system are discussed in sectio...', 'Result 1:\n14.l.4 High-Gain Antenna Acquisition Problem\n\nPrior to the television transmi...', 'Result 1:\n14.l.4 High-Gain Antenna Acquisition Problem\n\nPrior to the television transmi...'] +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:12:36 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:12:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first number in a list +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first number in list" +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first number in a list without knowing what the list refers to or its source? +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the meaning of "first" in a list +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first item in a list in general +2025-04-03 at 09:12:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:12:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the first numerical value in a common list +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "numbers in educational data" +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can you provide more context about what the list refers to or what the operations in Result 2 are applicable to? Where are these numbers coming from? +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 24.0 as the first number in a list +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: topic: lunar module table of contents first item in list +2025-04-03 at 09:12:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:12:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: describe common uses of arithmetic expressions +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What does the format and organization of the list in Result 1 suggest? +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: table of contents A.1 +2025-04-03 at 09:12:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 09:12:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:12:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical format for presenting data in Tables D-I and D-II? +2025-04-03 at 09:12:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 09:12:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical format for the numbers in Table D-I of the Command and Service Modules? +2025-04-03 at 09:12:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:12:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:12:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:12:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:12:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, False] +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:62 - Student lengths: [424, 622, 148, 163, 180, 174] +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:64 - Average student length: 285.17 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_correctness:66 - Length ratio: 47.53 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.542 ± 0.285 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.50 ± 1.50 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 2, 5, 2, 3, 0] +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: ++21.0 +3.0 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: ++21.0 +3.0 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 09:12:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\n+21.0 +3.0\n------\nResult 1:\n+21.0 +3.0\n------\nResult 2:\n-1.2 40.4 +0.4\n------\n...'] +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:12:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 09:12:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\n1.0 SUMMARY 1-1 2.0 INTRODUC...', 'Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\n(section ll.3).\n------\nResult 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitch holes with cracks extending from them on window shades +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracked stitch holes in window shades (by material or brand) +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracked stitch holes in fabric window shades +2025-04-03 at 09:12:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitch holes showing cracks in Aclar window shades +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracked stitch holes on window shades total +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracked stitch holes in fabric window shades +2025-04-03 at 09:12:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:12:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:13:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 80% of stitch holes crack rate in window shades +2025-04-03 at 09:13:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:13:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:13:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, True, True] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:62 - Student lengths: [491, 414, 582, 368, 351, 84] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:64 - Average student length: 381.67 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_correctness:66 - Length ratio: 38.17 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.333 ± 0.251 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 1.11 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 0, 2, 3, 2] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:13:01 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:13:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel consumption of specific service propulsion system in space exploration, e.g., STS-1, Apollo missions +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Space Shuttle main engine fuel consumption per thruster" +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel consumption service propulsion system amount +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service propulsion system fuel consumption for NASA sda lander or lunar mission +2025-04-03 at 09:13:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:13:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission service propulsion system fuel consumption +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Service propulsion system hydrazine consumption" +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Descent Propulsion System fuel consumption Apollo Mission +2025-04-03 at 09:13:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:13:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:13:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Service Propulsion System Fuel Consumption +2025-04-03 at 09:13:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: total amount of propellant used in Descent Propulsion System Apollo Mission +2025-04-03 at 09:13:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:13:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Service Descent Propulsion System Fuel Consumption +2025-04-03 at 09:13:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Descent Propulsion System fuel usage and total propellant used during lunar landing +2025-04-03 at 09:13:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command and Lunar Module Service Descent Propulsion System Fuel Consumption +2025-04-03 at 09:13:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Service Descent Propulsion System Fuel Consumption +2025-04-03 at 09:13:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Descent Propulsion System Fuel Consumption +2025-04-03 at 09:13:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:13:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:13:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, False, True] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1914, 497, 502, 28, 875, 531] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:64 - Average student length: 724.50 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_correctness:66 - Length ratio: 181.12 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.354 ± 0.284 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.83 ± 3.48 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [10, 2, 0, 1, 0, 4] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:13:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 09:13:25 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...'] +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:13:25 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:13:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:13:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Unplanned Minimum Impulse (UMI) engine firing investigation case studies +2025-04-03 at 09:13:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Unplanned Minimum Impulse engine firing at 32:21:49 investigation +2025-04-03 at 09:13:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 32:21:49 AGS engine firing, Tuesday [insert date here] time stamp +2025-04-03 at 09:13:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ + +2025-04-03 at 09:13:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 32:21:49 Transearth injection, digital autopilot error, roll manual attitude switch change +2025-04-03 at 09:13:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Unplanned minimum impulse engine firing, transition from rate-command to acceleration-command mode, 32:21:49 +2025-04-03 at 09:13:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:13:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:13:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, True, False] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:62 - Student lengths: [272, 932, 1021, 1776, 644, 405] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [112, 112, 112, 112, 112, 112] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:64 - Average student length: 841.67 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 112.00 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_correctness:66 - Length ratio: 7.51 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.079 ± 0.177 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 4.10 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 11, 0, 0] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ + +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 09:13:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:13:39 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:13:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:13:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the launch date and latitudinally/longitudinally centered seismic station that received the signal from the S-IVB impact? +2025-04-03 at 09:13:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:13:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:13:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:13:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, True] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:62 - Student lengths: [728, 1252, 1051, 1445, 464, 191] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:64 - Average student length: 855.17 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_correctness:66 - Length ratio: 85.52 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 0, 1] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:13:47 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:13:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Orbital insertion event delay" +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Artemis mission orbital insertion timing +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Orbital insertion time 12F5 Zuma Liferun mission schedule +2025-04-03 at 09:13:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 09:13:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:13:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 12F5 Zuma Liferun predicted re-entry time mission timeline +2025-04-03 at 09:13:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-03 at 09:13:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lifeguard mission first flight entry time +2025-04-03 at 09:13:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:13:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lifeguard mission orbital insertion time +2025-04-03 at 09:13:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:13:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:13:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Orbital insertion 12F5 Lifeguard first lunar orbit +2025-04-03 at 09:13:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:13:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection time 12F5 Lifeguard +2025-04-03 at 09:14:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:14:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Pericynthion arrival time Lifeguard 12F5 +2025-04-03 at 09:14:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:14:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Pericynthion arrival time and predicted entry time Lifeguard 12F5 +2025-04-03 at 09:14:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:14:05 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:14:05 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:14:05 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:62 - Student lengths: [329, 448, 373, 1368, 1397, 587] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:64 - Average student length: 750.33 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_correctness:66 - Length ratio: 150.07 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.196 ± 0.200 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 3.98 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 0, 0, 11, 0] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...'] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nManeuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft...'] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nManeuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft...', 'Result 1:\nThe prelaunch timeline was satisfactory, and adequate time was allotted for s...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nPotable water was obtained by periodically pressurizing the potable tank with...', 'Result 1:\nManeuver System Ignition time, hr:min:sec Firing time; sec Velocity change. f...', 'Result 1:\nManeuver System Ignition time, hr:min:sec Firing time; sec Velocity change. f...'] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 09:14:05 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:14:05 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:14:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:14:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon altitude unit +2025-04-03 at 09:14:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:14:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what unit is used to measure altitude above the lunar surface? +2025-04-03 at 09:14:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:14:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module altimeter unit +2025-04-03 at 09:14:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:14:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface altitude unit feet +2025-04-03 at 09:14:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:14:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module altimeter design or mechanics +2025-04-03 at 09:14:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 09:14:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon landing site altitude unit definition +2025-04-03 at 09:14:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:14:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module sun alignment unit +2025-04-03 at 09:14:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:14:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module sun/moon alignment measurement unit +2025-04-03 at 09:14:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:14:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module navigation altitude reference celestial body +2025-04-03 at 09:14:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:14:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module altitude calculation method +2025-04-03 at 09:14:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 09:14:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude reference lunar module +2025-04-03 at 09:14:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:14:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module altitude unit diameter +2025-04-03 at 09:14:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:14:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module altitude unit +2025-04-03 at 09:14:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:14:29 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:14:29 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:14:29 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, True] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1216, 588, 415, 884, 697, 853] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:64 - Average student length: 775.50 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_correctness:66 - Length ratio: 59.65 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.188 ± 0.270 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.33 ± 4.03 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 11, 0, 0, 3] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:14:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nd. The effectiveness of preflight crew trai...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nTo assure the alignment accuracy of the lun...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nTo assure the alignment accuracy of the lun...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nPerformance of Lunar Module Reaction Contro...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nPerformance of Lunar Module Reaction Contro...'] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:14:29 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:14:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature antenna heaters antenna off +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +the high-gain antenna switched from narrow beam to wide beam, because the panel, when separating, struck and damaged one of the antenna dishes. +------ + +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the normal operating temperature range for a satellite communications antenna and when to stop using antenna heaters? +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the normal operating temperature range for most antennas when they are turned off? +2025-04-03 at 09:14:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: results 1 antenna temperature minus 66 F wide beam antenna +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature monitoring system threshold for antenna heaters in a passive thermal control mode? +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the normal operating temperature range of an antenna when its heaters are turned off, and are the antennas designed or modified to withstand such low temperatures? +2025-04-03 at 09:14:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna operation wide beam mode minus 66 F +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What temperature reading are we looking for when the antenna heaters are switched to the passive mode? +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical operating temperature range for a steerable S-band antenna in passive mode when its heaters are turned off, in relation to the provided information about -66°F? +2025-04-03 at 09:14:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna wide beam manual mode minus 66 F +2025-04-03 at 09:14:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the standard operating temperatures for an S-band steerable antenna in various power transmission configurations, including low-power and receiving modes? +2025-04-03 at 09:14:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-gain antenna uplink and downlink signal strengths +2025-04-03 at 09:14:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:14:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the lowest operating temperature recorded for the S-band steerable antenna without the antenna heaters on? +2025-04-03 at 09:14:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna manual mode plus 23 degrees pitch plus 267 degrees yaw +2025-04-03 at 09:14:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:14:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What temperature range can a spacecraft antenna expect during onboard power off modes? +2025-04-03 at 09:14:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna manual mode minus 66 F wide beam signal strengths +2025-04-03 at 09:14:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can an S-band antenna be turned off and exposed to such low temperatures in a normal operation? +2025-04-03 at 09:14:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:14:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: scan limit warning function line minus 66 F +2025-04-03 at 09:14:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ +Result 2: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ + +2025-04-03 at 09:14:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature range of an antenna during low-power, low-bit-rate telemetry transmission using a 2lo-foot receiving antenna? +2025-04-03 at 09:14:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:14:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:14:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: scan limit function line point C plus 66 F +2025-04-03 at 09:14:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What temperature conditions are typically provided to 2lo-foot receiving antenna during low-power, low-bit-rate telemetry transmission in a space environment? +2025-04-03 at 09:14:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:14:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:14:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: scan limit function line point C minus 66 F +2025-04-03 at 09:14:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:14:57 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:14:57 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, False, False, False] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:62 - Student lengths: [978, 502, 201, 1861, 606, 618] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:64 - Average student length: 794.33 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 11.00 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_correctness:66 - Length ratio: 72.21 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.425 ± 0.437 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.67 ± 4.27 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [10, 0, 3, 9, 0, 0] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +the high-gain antenna switched from narrow beam to wide beam, because the panel, when separating, struck and damaged one of the antenna dishes. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ +Result 2: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ +Result 2: +Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated. + +mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 09:14:57 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:14:57 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:14:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first descent propulsion maneuver spacecraft average rise rate" +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: orbital maneuver Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the average ascent rate from liftoff for a typical commercial airliner or a glider? + +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: search "average rise rate to first descent phase spacecraft/aircraft" +2025-04-03 at 09:15:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:15:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:15:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Apollo 11 ascent to S-IVB separation average ascent rate + +2025-04-03 at 09:15:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:15:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Apollo 11 ascent average ascent rate to S-IVB separation + +2025-04-03 at 09:15:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:15:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Estimating the average ascent rate from liftoff to S-IVB separation + +2025-04-03 at 09:15:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:15:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Estimated ascents of vehicles such as the Saturn V + +2025-04-03 at 09:15:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:15:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Saturn V average ascent rate + +2025-04-03 at 09:15:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Apollo 12 average ascent rate during normal conditions + +2025-04-03 at 09:15:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:15:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:15:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:15:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, True, False, False] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:62 - Student lengths: [255, 924, 1906, 540, 371, 570] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:64 - Average student length: 761.00 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_correctness:66 - Length ratio: 76.10 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.254 ± 0.185 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 3.53 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 10, 0, 0, 1] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...'] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:15:22 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nboost phases were within acceptable spacecraft structural. design limits.\n---...'] +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:15:22 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:15:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:15:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system +2025-04-03 at 09:15:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system +2025-04-03 at 09:15:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:15:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide scrubber or filter +2025-04-03 at 09:15:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: planetary greenhouse effect lunar module liih efficiency +2025-04-03 at 09:15:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module emergency carbon dioxide scrubbing system +2025-04-03 at 09:15:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:15:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacesuit lithium hydroxide CO2 scrubber +2025-04-03 at 09:15:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:15:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacesuit with removable flow limiter +2025-04-03 at 09:15:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:15:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacesuit system modifications +2025-04-03 at 09:15:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:15:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacesuit system modifications +2025-04-03 at 09:15:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:15:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module carbon dioxide scrubbing system +2025-04-03 at 09:15:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:15:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:15:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:15:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1032, 642, 373, 419, 1225, 764] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:64 - Average student length: 742.50 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 28.00 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_correctness:66 - Length ratio: 26.52 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.352 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 3.30 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 1, 0, 0, 9] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:15:43 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:15:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of maximum lateral load experienced during S-IC boost compared to design value Saturn V Saturn V S-IC engine? +2025-04-03 at 09:15:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the percentage increase in maximum lateral loads during Intel S-IC boost compared to the design value? +2025-04-03 at 09:15:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 09:15:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:15:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Compare peak lateral loads during Saturn V S-IC boost to predicted design value and acceleration levels during peak S-IC boost oscillations +2025-04-03 at 09:15:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the maximum and peak lateral loads experienced during S-IC boost +2025-04-03 at 09:15:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:15:53 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:15:53 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:15:53 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, False] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:62 - Student lengths: [462, 114, 452, 430, 504, 319] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:64 - Average student length: 380.17 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_correctness:66 - Length ratio: 38.02 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.167 ± 0.251 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 1.11 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 0, 1, 0, 0, 0] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:15:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:15:53 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:15:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What sparked the development of flight menus on commercial airliners, and what contributes to their evolution? +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "airline flight menu development psychology" +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: psychology of menu design literature +2025-04-03 at 09:15:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:15:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: No evidence is found to indicate that the crew's preferences directly influenced the selection of flight menus. Can the menu selection be influenced by other factors such as cost, nutritional requirements, or regulatory constraints? +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "airline crew experience in food packaging and stowage" +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: packaging design for aerospace food systems +2025-04-03 at 09:16:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:16:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was the selection of flight menus for the Apollo missions influenced by NASA guidelines or regulations, such as nutritional requirements, food safety standards, or crew enhancement? +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission nutrition planning" +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA inspection protocols for aerospace food systems +2025-04-03 at 09:16:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were there any specific NASA guidelines or procedures that governed the menu selection, including factors such as calorie requirements, waste management, and crew-specific dietary needs? +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "weight and inertia values in spacecraft mass propertiesApollo lunar mission" +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission food rationing and distribution +2025-04-03 at 09:16:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What scientific basis or research informed the NASA recommendation for using juices over plain water during the abort phase to conserve water? +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 mission trajectory parameters" +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of food and beverages on spacecraft mass properties +2025-04-03 at 09:16:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the implications of water conservation efforts during the flight on crew performance and mission success? +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 trajectory analysis documentation" +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: impact of payload on spacecraft mass properties and payload capacity +2025-04-03 at 09:16:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:16:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How did the water usage strategy influence the crew's decision-making and performance during the mission, especially considering the available information on water conservation and limited availability? +2025-04-03 at 09:16:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 7 through Apollo 13 mission supplement documentation" +2025-04-03 at 09:16:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Could the water conservation efforts and crew's decision to consume juices as a preferred beverage have contributed to the mission's extended duration, and what implications does this have for future space missions? +2025-04-03 at 09:16:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:16:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 vs Apollo 12 supplement documentation key differences" +2025-04-03 at 09:16:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:16:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How did the crew's ability to conserve water and manage food resources contribute to the mission's overall success, particularly in relation to the available food options and the utility of the consumption of juices? +2025-04-03 at 09:16:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 documentation differences" +2025-04-03 at 09:16:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:16:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:16:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:16:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, True] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:62 - Student lengths: [557, 1653, 265, 2038, 518, 739] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [82, 82, 82, 82, 82, 82] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:64 - Average student length: 961.67 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 82.00 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_correctness:66 - Length ratio: 11.73 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.396 ± 0.426 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.17 ± 4.22 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 9, 0, 9, 0, 7] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:16:32 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:16:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the term for zero rotation of a shaft? +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero rotation or no rotation equivalent in mechanical systems +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero rotation output shaft equivalent +2025-04-03 at 09:16:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:16:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the term for zero optics in motor drive amplifiers +2025-04-03 at 09:16:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero output or no rotation equivalent in optical encoders or position sensors in servo systems +2025-04-03 at 09:16:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what does zero optics mode mean in motor drive amplifiers +2025-04-03 at 09:16:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode equivalent or reference position in optical encoders or position sensors +2025-04-03 at 09:16:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode reference position in servo systems +2025-04-03 at 09:16:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:16:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optic alignment in servo systems +2025-04-03 at 09:16:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:16:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:16:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:16:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:62 - Student lengths: [917, 978, 534, 512, 897, 550] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:64 - Average student length: 731.33 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_correctness:66 - Length ratio: 60.94 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.325 ± 0.368 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.50 ± 1.89 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 3, 0, 0, 5, 1] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:16:50 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:16:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:16:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of "report Performance Analysis" " +2025-04-03 at 09:16:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:16:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report performance analysis publication date +2025-04-03 at 09:16:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:16:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Performance Analysis report from NASA Apollo mission 1960s or 1970s" " +2025-04-03 at 09:16:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:16:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission reports performance analysis +2025-04-03 at 09:16:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:16:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:16:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Performance Analysis or Apollo 7 Performance Analysis" " +2025-04-03 at 09:16:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:16:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance of Command and Service Module December 1969 +2025-04-03 at 09:16:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:16:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Performance Analysis report October 1969" " +2025-04-03 at 09:17:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:17:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Reports Performance Analysis sections +2025-04-03 at 09:17:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 7 Performance Analysis report" " +2025-04-03 at 09:17:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:17:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: original publication dates of Apollo mission analysis September 1969 or 1970" " +2025-04-03 at 09:17:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:17:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:17:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 Performance Analysis report March 1970" " +2025-04-03 at 09:17:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:17:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:17:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:17:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:17:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:62 - Student lengths: [548, 184, 339, 301, 393, 478] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:64 - Average student length: 373.83 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_correctness:66 - Length ratio: 28.76 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.292 ± 0.415 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 2.73 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 7, 4, 0, 0] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:17:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 09:17:08 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:17:08 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:17:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:17:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: International Space Exploration Mission Evaluation Team approved document year +2025-04-03 at 09:17:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:17:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:17:13 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:17:13 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:17:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, False, False] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:62 - Student lengths: [383, 28, 401, 510, 597, 611] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:64 - Average student length: 421.67 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_correctness:66 - Length ratio: 105.42 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.058 ± 0.130 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.17 ± 0.37 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 0, 0, 0] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:17:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:17:14 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:17:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa center location +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of NASA Johnson Space Center +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa manned spacecraft center location +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "nasa manned spaceflight management center location" +2025-04-03 at 09:17:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:17:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Manned Spacecraft Center facilities locations" +2025-04-03 at 09:17:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:17:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:17:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:17:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:17:21 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:62 - Student lengths: [365, 525, 234, 203, 312, 652] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:64 - Average student length: 381.83 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 20.00 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_correctness:66 - Length ratio: 19.09 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.258 ± 0.190 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.83 ± 0.69 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 1, 1, 0, 2] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:17:21 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:17:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:17:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was pilot involved descriptive function during NASA Apollo mission +2025-04-03 at 09:17:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:17:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Pilot's role in describing spacecraft systems during a mission: relevance to mission progress +2025-04-03 at 09:17:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 09:17:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 midcourse correction navigation difficulties with television broadcast +2025-04-03 at 09:17:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 09:17:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were crew members involved in motion control and system activation during the lunar module's midcourse correction? +2025-04-03 at 09:17:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:17:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 navigation during post-pericynthion phase +2025-04-03 at 09:17:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were astronauts involved in describing critical mission maneuvers involving the lunar module? +2025-04-03 at 09:17:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 vs Apollo 13 navigation and guidance system comparison +2025-04-03 at 09:17:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:17:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did astronauts describe maneuvers to avoid gimbal lock during lunar module undocking? +2025-04-03 at 09:17:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:17:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Pilot role during critical phases of Apollo 13 +2025-04-03 at 09:17:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the lunar module pilot participate in establishing entry procedures for the mission? +2025-04-03 at 09:17:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew response to emergency during critical phases +2025-04-03 at 09:17:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:17:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the lunar module pilot execute critical maneuvers, including coalignment and alignment operations during entry? +2025-04-03 at 09:17:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 09:17:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:17:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew training skills and emergency response +2025-04-03 at 09:17:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:17:49 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:17:49 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:17:49 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, True, True] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:62 - Student lengths: [387, 2032, 740, 521, 568, 352] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:64 - Average student length: 766.67 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 3.00 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_correctness:66 - Length ratio: 255.56 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.237 ± 0.363 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.83 ± 4.10 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 7, 10, 0, 0, 0] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe television presentation during the midcourse correction maneuver, as well...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe performance of the command and service module systems is discussed in thi...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...'] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 09:17:49 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:17:49 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:17:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: data center cooling capacity calculation water weight +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +------ + +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "reduced power condition in cooling systems" +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What does "reduced power condition" refer to in the context of cooling and battery performance? +2025-04-03 at 09:17:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:17:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:18:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:18:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, False, True] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:62 - Student lengths: [569, 642, 1069, 327, 865, 945] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:64 - Average student length: 736.17 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_correctness:66 - Length ratio: 92.02 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.175 ± 0.175 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.50 ± 0.50 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 1, 1, 1] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +------ + +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:18:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...'] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking·is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:18:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...'] +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:18:02 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:18:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: search "nickel content in spacecraft or planetary lander command module or hot water port" +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port Apollo CSM +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module Nikita Kramchenko hot water port nickel content +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: orbital station 3 > S made by Apollo's Command Module +2025-04-03 at 09:18:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:18:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-containing battery type +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in Command Module hot water port Apollo mission +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo l3 command module +2025-04-03 at 09:18:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:18:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery internal chemistry nickel +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo hot water port nickel content and electroplating +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: manual Apollo l3 hot water compartment +2025-04-03 at 09:18:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in lithium-ion battery insulation +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module hot water port nickel content and Apollo mission +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service Module modifications +2025-04-03 at 09:18:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:18:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickelode thermal interface material +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 hot water port nickel content +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft heat exchanger system +2025-04-03 at 09:18:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:18:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel carrier in thermal interface materials +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 command module fluids testing nickel contamination +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module heat exchanger system +2025-04-03 at 09:18:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:18:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel-based nickel/aluminum alloy +2025-04-03 at 09:18:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:18:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module systems interaction with hot water port +2025-04-03 at 09:18:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:18:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:18:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel plating of transducer components +2025-04-03 at 09:18:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:18:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:18:28 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:18:28 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:18:28 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, False] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:62 - Student lengths: [509, 647, 312, 502, 382, 371] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:64 - Average student length: 453.83 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 20.00 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_correctness:66 - Length ratio: 22.69 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.454 ± 0.391 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 6.17 ± 7.71 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 8, 0, 0, 6, 22] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:18:28 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe experiment consisted of an electronics package and two sensor probes whic...', 'Result 1:\nThe experiment consisted of an electronics package and two sensor probes whic...', 'Result 1:\nThe transducer is a variable reluctance instrument actuated by differential p...', 'Result 1:\nThe transducer is a variable reluctance instrument actuated by differential p...'] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:18:28 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nThe structure in the forward...', 'Result 1:\nThe experiment consisted of an electronics package and two sensor probes whic...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe performance of the command and service module systems is discussed in thi...'] +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:18:28 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:18:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:18:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of resolvers used for shaft axis +2025-04-03 at 09:18:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:18:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resolver used only for shaft axis +2025-04-03 at 09:18:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 09:18:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:18:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver slip ring vacuum testing +2025-04-03 at 09:18:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 09:18:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:18:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver slip ring wipes axis system +2025-04-03 at 09:18:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:18:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:18:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:18:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:18:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:62 - Student lengths: [555, 560, 460, 595, 982, 420] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [19, 19, 19, 19, 19, 19] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:64 - Average student length: 595.33 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 19.00 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_correctness:66 - Length ratio: 31.33 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.167 ± 0.251 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.67 ± 1.11 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 3, 1, 0, 0, 0] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 09:18:39 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:18:39 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:18:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze in water and in air? +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to impurities in a frozen state? +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Frost formation on contaminants in surface finish inspection + + +2025-04-03 at 09:18:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:18:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:18:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How do contaminants freeze in a gas or liquid state? +2025-04-03 at 09:18:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:18:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Material properties of solder flux, glass beads, and fibers under freezing conditions. + + +2025-04-03 at 09:18:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:18:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:18:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how does a liquid become a gas? +2025-04-03 at 09:18:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:18:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the presence of moisture in a vacuum system affect the freezing point of the contaminants? + + +2025-04-03 at 09:18:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:18:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:18:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does a liquid turn into a gas? +2025-04-03 at 09:18:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:18:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How do frozen contaminants in a vacuum system behave when the temperature is increased? + + +2025-04-03 at 09:18:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:18:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does a gas condense and a liquid freeze? +2025-04-03 at 09:19:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the vaporization temperature of a contaminant affect the pressure rise rate in a vacuum system? + + +2025-04-03 at 09:19:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the relationship between contaminant freezing and condensation? +2025-04-03 at 09:19:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the vapor pressure of hydrogen in its liquid and vapor states at typical space temperatures? + + +2025-04-03 at 09:19:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:19:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what happens when a gas freezes? +2025-04-03 at 09:19:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical pressure range for a helium tank during a lunar landing? + + +2025-04-03 at 09:19:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:19:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the difference between freezing and condensation? +2025-04-03 at 09:19:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can a vapor condense to form a solid? +2025-04-03 at 09:19:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the relationship between solid formation and thermal conductivity? +2025-04-03 at 09:19:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:19:16 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:19:16 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:19:16 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:62 - Student lengths: [156, 433, 263, 1593, 987, 1717] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [92, 92, 92, 92, 92, 92] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:64 - Average student length: 858.17 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 92.00 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_correctness:66 - Length ratio: 9.33 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.292 ± 0.347 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.50 ± 4.07 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 3, 10, 0, 8] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...'] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:19:16 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:19:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission launch date +2025-04-03 at 09:19:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission launch date +2025-04-03 at 09:19:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission launch date +2025-04-03 at 09:19:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 surface mission launch date +2025-04-03 at 09:19:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 saft stage test launch +2025-04-03 at 09:19:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 09:19:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 saft stage launch date +2025-04-03 at 09:19:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:19:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 saft stage launch date +2025-04-03 at 09:19:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:19:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 wikipedia +2025-04-03 at 09:19:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:19:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:19:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:19:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, True] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:62 - Student lengths: [521, 425, 2018, 973, 345, 527] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:64 - Average student length: 801.50 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_correctness:66 - Length ratio: 66.79 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.352 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.33 ± 2.56 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 7, 0, 0, 0] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11��1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 09:19:32 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:19:32 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:19:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Evolution of space and atmospheric imaging 1800s- present +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: When did the NASA Earth Weather Camera (HEVEC) launch and end? +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is there a well-documented record or historical timeline of Earth weather photography from reputable sources? +2025-04-03 at 09:19:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:19:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the name of the specific Apollo mission that included the Earth Weather Photography experiment? +2025-04-03 at 09:19:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:19:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: History of operational Earth weather photography in space exploration programs, including duration and significance +2025-04-03 at 09:19:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the mission number of the Apollo spaceflight that included the Earth Weather Photography experiment? +2025-04-03 at 09:19:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:19:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How common were Earth weather photography operations in space exploration programs and were there any standard durations established for such operations? +2025-04-03 at 09:19:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the launch and landing dates of Apollo 12? +2025-04-03 at 09:19:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:19:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Standard operating practices for Earth weather observation missions and duration of such observations +2025-04-03 at 09:19:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:19:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ground-based Earth weather observation missions, such as weather satellites, or mountain weather stations and their projected operational times +2025-04-03 at 09:19:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is stereographic photogrammetry used for cloud height determination in operational weather observation missions? If yes, what is the typical altitude range and duration of these operations? +2025-04-03 at 09:19:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the Advanced Technology Satellites I and IIl still in operation? If yes, how long have they been operational? +2025-04-03 at 09:19:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Advanced Technology Satellite I was out of operation on the day of photography. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:19:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is there any information on the operational history of geosynchronous weather satellites? If yes, how long have they been in operation? +2025-04-03 at 09:19:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ + +2025-04-03 at 09:19:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did geosynclastic Earth weather satellites become a standard operational platform for weather observation? If yes, how long have they been in operation? +2025-04-03 at 09:19:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:19:59 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:19:59 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:19:59 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, False, False, True] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:62 - Student lengths: [980, 249, 209, 851, 1029, 861] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [21, 21, 21, 21, 21, 21] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:64 - Average student length: 696.50 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 21.00 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_correctness:66 - Length ratio: 33.17 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.279 ± 0.347 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 3.35 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 0, 0, 5, 9] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Advanced Technology Satellite I was out of operation on the day of photography. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:19:59 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:20:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query "NASA Lyndon B. Johnson Space Center location +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center location +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: National Aeronautics and Space Administration Lyndon B. Johnson Space Center location +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-03 at 09:20:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:20:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: exact address of NASA Manned Spacecraft Center +2025-04-03 at 09:20:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:20:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: has the NASA Manned Spacecraft Center moved +2025-04-03 at 09:20:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:20:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:20:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:62 - Student lengths: [173, 530, 542, 425, 739, 72] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:64 - Average student length: 413.50 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 14.00 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_correctness:66 - Length ratio: 29.54 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.342 ± 0.188 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.17 ± 0.90 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 3, 1, 1, 0, 1] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:20:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPOSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Mann...'] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 09:20:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:20:09 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:20:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations peak amplitude +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillations peak amplitude 33.7g when?" + +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +when did the S-II luminosity experiment begin luminosity measurements at the LHC? + +2025-04-03 at 09:20:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:20:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:20:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation peak amplitude 33.7g Apollo 13" + +2025-04-03 at 09:20:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +details of s-II vibration S-II second stage details S-II crossbeam oscillations lunar mission events + +2025-04-03 at 09:20:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 09:20:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:20:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation peak amplitude 33.7g Apollo 13 S-IVB" + +2025-04-03 at 09:20:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation 33.7g amplitude during Apollo 13 ascent 52 minutes" + +2025-04-03 at 09:20:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation 33.7g amplitude Apollo 13 staging S-II to S-IVB" + +2025-04-03 at 09:20:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 09:20:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation 33.7g amplitude during S-IVB ignition Apollo 13" + +2025-04-03 at 09:20:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 09:20:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation 33.7g amplitude during S-II boost Apollo 13 330.6 seconds" + +2025-04-03 at 09:20:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-II crossbeam oscillation 33.7g amplitude 330.6 seconds after liftoff" +2025-04-03 at 09:20:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:20:31 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:20:31 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:20:31 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, True] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:62 - Student lengths: [391, 523, 2047, 1270, 1064, 517] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:64 - Average student length: 968.67 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_correctness:66 - Length ratio: 74.51 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.300 ± 0.350 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 2.85 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 1, 8, 0, 2, 0] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 09:20:31 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nIgnition and lift-off occurred on schedule. A listing_of major flight plan ev...', 'Result 1:\nIgnition and lift-off occurred on schedule. A listing_of major flight plan ev...'] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 09:20:31 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:20:31 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:20:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"What are the adverse effects of extraneous discharges or faults in electrical circuits, specifically in experiments with figure 5.6-1?" + + +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "adverse effects of extraneous firing in high-energy applications" +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what are the adverse effects of extraneous firings in electrical circuits? +2025-04-03 at 09:20:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:20:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA-5-70-5825 adverse effects of extraneous firings Apollo mission" +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of extraneous firings on spacecraft navigation and alignment +2025-04-03 at 09:20:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:20:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 post-flight testing and investigation" +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: impact of transient electrical events on spacecraft attitude control during thermal maneuvers +2025-04-03 at 09:20:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:20:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 extraneous firing investigation" +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of transient electrical events on attitude control systems and thermal regime during spacecraft transit phases +2025-04-03 at 09:20:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:20:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 mission risk assessment and extraneous firing mitigation" +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: impact of transient electrical events on spacecraft attitude control, power management, and gyroscopic systems during lunar transit phases +2025-04-03 at 09:20:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:20:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 abort profile and emergency system configurations" +2025-04-03 at 09:21:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 09:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: root cause of dc bus voltage dip and short circuit in lunar module descent stage during ascent +2025-04-03 at 09:21:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:21:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 trajectory recomputation and extraneous firings" +2025-04-03 at 09:21:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:21:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: causes of dc circuit faults during recurrent electrical transients during lunar module power system operation +2025-04-03 at 09:21:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:21:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:21:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 navigation and guidance" +2025-04-03 at 09:21:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:21:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:21:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:21:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, True] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:62 - Student lengths: [427, 1056, 28, 605, 1906, 2031] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [41, 41, 41, 41, 41, 41] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:64 - Average student length: 1008.83 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 41.00 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_correctness:66 - Length ratio: 24.61 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.354 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.17 ± 4.41 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 6, 0, 0, 8, 11] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:21:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:21:09 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:21:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IVB engine cutoff time" or "Apollo S-IVB engine shutdown time" +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the time of S-IVB mission re-entry burn for Space Shuttle program +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the launch time of S-IVB engine cutoff for a space probe? +2025-04-03 at 09:21:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo S-IVB engine shutdown time" or "Apollo S-IVB cutoff time, Saturn V" +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apost11 s-ivb engine cutoff time +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 S-IVB engine cutoff time +2025-04-03 at 09:21:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:21:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Related/Apollo l2 l2 s-ivb engine cutoff +2025-04-03 at 09:21:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 S-IVB engine cutoff parking orbit detail +2025-04-03 at 09:21:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:21:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ap11 s-ivb engine cutoff deviations +2025-04-03 at 09:21:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff mission context +2025-04-03 at 09:21:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 09:21:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: result 1 s-ivb engine cutoff converted to decimal time +2025-04-03 at 09:21:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch sequence and flight times +2025-04-03 at 09:21:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:21:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apoll l2 vs. ap11 s-ivb engine cutoff timing +2025-04-03 at 09:21:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff S-IVB launch Apollo 9 +2025-04-03 at 09:21:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apoll l2 sivborbital insertion duration +2025-04-03 at 09:21:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff initial +2025-04-03 at 09:21:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ap11 mission timeline activities with 00:12:30 +2025-04-03 at 09:21:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB ignition time +2025-04-03 at 09:21:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:21:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB S 4 mission launch sequence +2025-04-03 at 09:21:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:34 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:21:34 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:21:35 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, True, True, True] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:62 - Student lengths: [304, 334, 233, 1919, 1326, 392] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:64 - Average student length: 751.33 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_correctness:66 - Length ratio: 93.92 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.400 ± 0.427 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.17 ± 3.85 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 2, 8, 9, 0] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 09:21:35 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:21:35 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:21:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission reports +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission media reports +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission reports +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 8 mission reports" +2025-04-03 at 09:21:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:21:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:21:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission total number of reports +2025-04-03 at 09:21:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:21:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission report number 1, 2, 3, 6 +2025-04-03 at 09:21:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:21:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission comprehensive report list +2025-04-03 at 09:21:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:21:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission report count +2025-04-03 at 09:21:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:21:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission supplemental report list +2025-04-03 at 09:21:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:21:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 supplemental report numbers +2025-04-03 at 09:21:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:21:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 total number of supplemental reports +2025-04-03 at 09:21:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:21:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission report count +2025-04-03 at 09:21:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:21:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:21:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 report list information +2025-04-03 at 09:21:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:21:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 mission report related to Apollo 8 +2025-04-03 at 09:21:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:21:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 and Apollo 10 supplemental report overlap +2025-04-03 at 09:22:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:22:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 mission report supplement, series , page 36 +2025-04-03 at 09:22:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 total number of supplemental reports +2025-04-03 at 09:22:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:22:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:22:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:22:03 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, True, True] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1329, 1134, 277, 269, 2049, 832] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:64 - Average student length: 981.67 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 1.00 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_correctness:66 - Length ratio: 981.67 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.433 ± 0.392 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.83 ± 3.34 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 0, 1, 0, 7, 1] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:22:03 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:22:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "IIT Kanpur Placement SC-009 Supercircular location" +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: asx sc 009 supercircular february 26 1966 location +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Supercraft AS-201 SC-009 location February 26 1966 +2025-04-03 at 09:22:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:22:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 entry test zero-G location February 1966" +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo program launch site 1966 +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 flight April 1966 +2025-04-03 at 09:22:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:22:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 or Apollo 10 entry location February 1966" +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa missiop launch site 1966 +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module lunar surface recovery +2025-04-03 at 09:22:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:22:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo Superground Circular location and date" +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: white sands missile range launch +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar surface AS-201 SC-009 +2025-04-03 at 09:22:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:22:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:22:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 supercritical helium pressure +2025-04-03 at 09:22:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:22:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 helium pressure +2025-04-03 at 09:22:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:22:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 AS-201 SC-009 no, S-IVB Y to S-IVB Y comparison +2025-04-03 at 09:22:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:26 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:22:26 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:22:26 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:62 - Student lengths: [467, 471, 874, 273, 42, 1898] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:64 - Average student length: 670.83 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 7.00 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_correctness:66 - Length ratio: 95.83 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.292 ± 0.311 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.83 ± 3.08 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 5, 0, 4, 8] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:22:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ["Result 1:\nThe Iwo Jima's position was established accurately using a satellite navigati...", 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:22:26 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:22:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:22:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB vibration duration +2025-04-03 at 09:22:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB mission duration vibration +2025-04-03 at 09:22:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: periods of S-IVB vibration S-IVB mission +2025-04-03 at 09:22:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB translunar injection vibration hours +2025-04-03 at 09:22:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB re-entry vibration duration +2025-04-03 at 09:22:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration periods +2025-04-03 at 09:22:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB re-entry vibration correction +2025-04-03 at 09:22:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration periods +2025-04-03 at 09:22:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration event duration re-entry +2025-04-03 at 09:22:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: range of time periods when S-IVB vibration occurred +2025-04-03 at 09:22:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:22:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of S-IVB vibration during Transposition and Docking +2025-04-03 at 09:22:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:22:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: unique start time of S-IVB vibration during Transposition and Docking +2025-04-03 at 09:22:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration start time and duration +2025-04-03 at 09:22:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:22:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of S-IVB vibration +2025-04-03 at 09:22:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:22:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:22:48 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:62 - Student lengths: [234, 403, 368, 422, 540, 1336] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:64 - Average student length: 550.50 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 15.00 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_correctness:66 - Length ratio: 36.70 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.448 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.33 ± 3.50 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 0, 0, 0, 0, 9] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:22:48 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:22:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic layer materials for structural integrity +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material was used for the insulation between the two shells of a liquid helium tank? +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: insulation material for helium tank shell +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Based on this hint about cryogenic applications, I'm going to phrase my next search query +2025-04-03 at 09:22:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:22:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: contamination stainless steel cryogenic tank +2025-04-03 at 09:23:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Given the previous information about the oxygen tank and the Apollo mission, I'll try a more general search +2025-04-03 at 09:23:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 09:23:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reactive materials oxygen tank burning +2025-04-03 at 09:23:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Since I couldn't find a specific answer to the previous question, I will try a different search query +2025-04-03 at 09:23:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-03 at 09:23:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fire hazard materials oxygen tank +2025-04-03 at 09:23:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:23:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal expansion stainless steel cryogenic +2025-04-03 at 09:23:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:23:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: stainless steel vs other materials cryogenic +2025-04-03 at 09:23:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:23:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:23:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:23:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, True, False] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:62 - Student lengths: [931, 1932, 150, 521, 435, 333] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:64 - Average student length: 717.00 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_correctness:66 - Length ratio: 44.81 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.288 ± 0.237 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.33 ± 4.89 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 14, 1, 0, 2, 3] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:23:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nCryogenic storage system operation was satisfactory until 46:40:09, when the ...', 'Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...'] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(section ll.3). +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-03 at 09:23:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nBecause of a sudden loss of pressure at approximately 56 hours from one of th...', 'Result 1:\n(section ll.3).\n------\nResult 2:\nPotable water was obtained by periodically p...'] +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:23:14 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:23:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the initial pitch angle of a spacecraft typically set for launch. +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the pitch of the SpaceX Starship launch vehicle's initial stage? +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the pitch of the initial launch vehicle of a spacecraft? +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: What is the initial pitch of the X-43A? +2025-04-03 at 09:23:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Initial pitch maneuver data for a lunar landing mission. +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the initial pitch of the Saturn V rocket? +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the required pitch attitude for initial launch and docking phases of a NASA spacecraft? +2025-04-03 at 09:23:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:23:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: umanual pitch maneuver Apollo lunar module. +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V initial pitch electrostatic potential +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft docking requirements specification +2025-04-03 at 09:23:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 09:23:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module initial pitch maneuver angle. +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch vehicle initial pitch +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft pitch angle for initial separation, docking, and trans-Earth injection +2025-04-03 at 09:23:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:23:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V pitch at liftoff +2025-04-03 at 09:23:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft pitch and roll attitudes for initial separation and docking +2025-04-03 at 09:23:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V pitch at lift-off radio noise +2025-04-03 at 09:23:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:23:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch timing corona discharges +2025-04-03 at 09:23:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 09:23:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:23:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch phase duration +2025-04-03 at 09:23:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:23:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:23:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:23:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, False, True, True] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:62 - Student lengths: [746, 1849, 797, 704, 909, 851] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:64 - Average student length: 976.00 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_correctness:66 - Length ratio: 244.00 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.508 ± 0.412 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.00 ± 2.94 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [4, 8, 0, 0, 5, 1] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:23:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...'] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:38 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...'] +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:23:38 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:23:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: arc fault circuit interrupter contents +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What component escapes from the left-hand electrical circuit interrupter +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What hazardous material is commonly described to have 'escaped' in electrical circuit interrupter and linked to shocks or fatalities? +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Electrical circuit interrupter components that can escape +2025-04-03 at 09:23:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:23:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical circuit interrupter command module space shuttle escaped gas +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical circuit interrupter gas leak space mission +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What caused a gas supercritical pressure rise in a lunar module's helium system during an Apollo mission? +2025-04-03 at 09:23:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:23:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of power system is used by space mission interrupters +2025-04-03 at 09:23:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What gas escaped from the pressure rise in the Apollo descent stage's helium tank during testing? +2025-04-03 at 09:23:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:23:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what causes electrical system gas leaks in space missions +2025-04-03 at 09:23:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-03 at 09:23:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the issue with the Apollo 13 helium tank pressure rise during the initial cold-soak period on mission? +2025-04-03 at 09:23:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:23:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: prevention of gas leaks in electrical circuit breaker components +2025-04-03 at 09:23:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What caused the initial helium pressure increase in the lunar module's descent propulsion system during the Apollo 13 mission? +2025-04-03 at 09:23:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:23:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:23:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the cause of the abnormally high helium pressure increase prior to the Apollo 13 launch? +2025-04-03 at 09:23:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:23:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:23:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What caused the abnormally high helium pressure prior to the Apollo 13 lunar module launch? +2025-04-03 at 09:23:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:23:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Compare and contrast the two instances of abnormally high helium pressure increases in the Apollo 13 mission and the Apollo 13 countdown demonstration test. +2025-04-03 at 09:24:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:24:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:24:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, True] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:62 - Student lengths: [534, 725, 1717, 628, 488, 901] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:64 - Average student length: 832.17 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 14.00 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_correctness:66 - Length ratio: 59.44 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.458 ± 0.391 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.67 ± 2.92 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 5, 8, 1, 0, 0] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:24:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nBased upon an analysis of the interrupter design, its location, and its relat...'] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 09:24:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:24:01 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:24:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What pressure range did the helium tank's rise-rate characteristics increase during the Apollo 13 mission's countdown demonstration test? +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: he helium tank rise-rate characteristics use Montgomery Ward 7-25 and pressure +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 helium tank pressure rise rate during test +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure rise rate Apollo 13 countdown demonstration test +2025-04-03 at 09:24:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:24:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What pressure range of helium saturation does the Apollo 13 mission's countdown demonstration test indicate as the threshold for abnormal rise-rate characteristics, leading to a helium tank pressure rise of 14.9 psi/hour? +2025-04-03 at 09:24:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range helium tank pressure rise rate Apollo 13 countdown demonstration test 640-900 psi +2025-04-03 at 09:24:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the specific pressure lower bound at which the helium tank's rise-rate characteristic begins to deviate abnormally, according to the Apollo 13 countdown demonstration test results? +2025-04-03 at 09:24:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure rise rate Apollo 13 countdown demonstration test normal range 640-900 psi +2025-04-03 at 09:24:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What pressure range of helium saturation is the threshold for the abnormal rise-rate characteristic to be around 14.9 psi/hour, as indicated by the heat-leak test during the Apollo 13 countdown demonstration test? +2025-04-03 at 09:24:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure rise rate anomaly helium tank Apollo 13 ascent < 600 psi +2025-04-03 at 09:24:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the specific pressure lower bound of helium saturation at which the rise-rate characteristic becomes abnormal, exactly or to some extent, where helium's rise-rate increased by l4.9 psi/hour, taken from the 900 psia onwards. +2025-04-03 at 09:24:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature helium pressure 640 psia Apollo 13 countdown demonstration test +2025-04-03 at 09:24:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the pressure at which the helium's rise-rate increases abnormally, exactly, during the Apollo 13 countdown demonstration test or on the flight conditions, to a rate of 14 psi/hour? Alternatively, conservatively answer what is the pressure at which the helium tank's rise-rate increased from normal to abnormal. +2025-04-03 at 09:24:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:23 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:24:23 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:24:23 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:62 - Student lengths: [966, 152, 252, 820, 838, 672] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:64 - Average student length: 616.67 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 24.00 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_correctness:66 - Length ratio: 25.69 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.433 ± 0.392 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 2.41 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 6, 1, 1, 5, 0] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 09:24:23 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:24:23 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:24:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:24:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is throttle position in a vehicle control system and what does it control? +2025-04-03 at 09:24:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:24:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the final throttle position of the ill-fated 1986 Hindenburg airship? +2025-04-03 at 09:24:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:24:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the final throttle position during the descent propulsion system maneuver to a free-return trajectory? +2025-04-03 at 09:24:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:24:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the final throttle position of the hydrogen fuel cells in the Space Shuttle's main propulsion system? +2025-04-03 at 09:24:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 09:24:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the final throttle position during the transearth injection maneuver? +2025-04-03 at 09:24:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the final thrust or throttle position of a Space Shuttle main engine launch? +2025-04-03 at 09:24:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the final throttle position during the transearth injection maneuver? +2025-04-03 at 09:24:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the final thrust or throttle values of the Space Shuttle main engines during the transearth injection maneuver? +2025-04-03 at 09:24:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the final throttle position of the transearth injection maneuver? +2025-04-03 at 09:24:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the minimum throttle position of the Space Shuttle's descent propulsion system? +2025-04-03 at 09:24:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:40 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:24:40 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:24:40 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, False, False] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1043, 34, 1044, 1198, 549, 274] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:64 - Average student length: 690.33 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 13.00 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_correctness:66 - Length ratio: 53.10 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.317 ± 0.448 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.67 ± 2.36 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 5, 0, 0, 0, 5] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:24:40 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nDuring the mission, the fuel cells supplied approximately l20 kW-h of energy ...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:24:40 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:24:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:24:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training section document +2025-04-03 at 09:24:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 09:24:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training operations handbook +2025-04-03 at 09:24:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-03 at 09:24:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:24:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training program outline section 8.0 +2025-04-03 at 09:24:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 09:24:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: PIOTS training manual +2025-04-03 at 09:24:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:24:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Astronaut training command module operations +2025-04-03 at 09:24:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:24:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot astronaut training lunar module +2025-04-03 at 09:24:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:24:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pilot training procedures +2025-04-03 at 09:24:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:24:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training module simulator +2025-04-03 at 09:24:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 09:24:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: autopilot training for pilots +2025-04-03 at 09:24:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 09:24:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flight director training +2025-04-03 at 09:24:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:24:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:24:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo 8.0 training +2025-04-03 at 09:24:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:24:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:24:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:24:56 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, True] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:62 - Student lengths: [535, 401, 420, 280, 344, 417] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:64 - Average student length: 399.50 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 3.00 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_correctness:66 - Length ratio: 133.17 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.217 ± 0.352 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 3.67 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 0, 0, 0, 10, 0] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...'] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nCommander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., an...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nThe first midcourse correction maneuver, performed at the second option point...', 'Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\nCommander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., an...'] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +2025-04-03 at 09:24:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:24:56 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:24:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft isolation valve testing simulation" or "Launch and operational stress testing for isolation valves in spacecraft" +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vibration testing for isolation valves in space/ +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum chamber testing isolation valve verification +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft electrical isolation valve testing methods radiation hardness +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "validation testing for isolation valves on spacecraft" +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft isolation valve testing procedures" +2025-04-03 at 09:25:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft fluid control valve wiring verification" or "Isolation valve wiring validation" +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: wiring harness isolation valve verification resistance testing" +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft wiring validation methods component validation testing procedures +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "wiring resistance testing of isolation valves for reaction control system in spacecraft" +2025-04-03 at 09:25:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve walk test" or "Voltage drop testing of isolation valves" +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation testing isolation valves wired correctly +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module electrical isolation valve verification testing +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "circuit verification wiring harness procedure" or "spacelist validation resistance testing method" +2025-04-03 at 09:25:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve functional testing" or "Verification of isolated valve operation with and without power" +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: miswired isolation valve circuit validation testing reverse polarity +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inductive coupling isolation valve testing reverse-polarity voltage closure +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "wiring airtightness testing" + +It was useless as the prior test found that functional checks are insufficient for verifying wiring. + +Let me provide a new search query to see if the answer is available "circuit verification methods for spacecraft isolation valves" +2025-04-03 at 09:25:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve testing without power" or "Testing for non-functional operation" +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of isolation valve wiring scheme +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance testing isolation valve verification +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "inductive coupling based isolation valve control" "valve wiring validation through inductive coupling" +2025-04-03 at 09:25:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Verification of isolation valve operation using functional checks only" or "Testing isolation valves with functional checks" +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of isolation valve functionality testing resistance checks +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation testing for isolation valves +2025-04-03 at 09:25:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Resistive testing for proper isolation valve wiring" or "Verification of isolation valve wiring through resistance checks" +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of isolation valve circuit wiring +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical insulation of isolation valves +2025-04-03 at 09:25:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:25:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:25:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of isolation valve validation +2025-04-03 at 09:25:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: in-circuit test isolation valves +2025-04-03 at 09:25:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:25:30 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:25:30 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, True] +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1922, 424, 1922, 1909, 543, 1024] +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:64 - Average student length: 1290.67 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 17.00 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_correctness:66 - Length ratio: 75.92 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.671 ± 0.282 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.33 ± 3.09 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [7, 1, 8, 8, 1, 7] +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 6.0 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 1.000 +2025-04-03 at 09:25:30 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:25:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what are the components of an automotive engine coolant system + + +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of the original sensing point for the water separator drain tank in a typical automotive engine? +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pump design onboard marine vessel sensing point water separator drain tank location +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "marine engine water separator sensor location" +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Original location of sensing point in water separator drain tank pump system" +2025-04-03 at 09:25:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:25:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: automotive water separator location in engine coolant system + + +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of sensing point for the water separator drain tank in the Space Shuttle's ascent oxygen system +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo suit pressure transducer location +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 15 life support system water separator drain tank sensing point" +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "skylab life support system sensors location" +2025-04-03 at 09:25:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:25:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of water separator drain tank in lunar module + + +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location and type of valve or component sensing the water separator drain on ascent oxygen tank 2 +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure transducer location for water separator tank moon landing +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 15 mission water separator drain tank location" +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "skylab water separator drain tank location" +2025-04-03 at 09:25:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:25:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of water separator drain tank in command module + + +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of shutoff valve in ascent oxygen tank 2, nearing or at manifold 14-l9 +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank location moon landing Apollo +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 15 mission specifics fossil boring waste water separator tank location" +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module water separator drain tank location" +2025-04-03 at 09:25:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:25:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of water separator drain tank in service module + + +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 water separator drain tank location +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 abor thermal water separator drain tank location" +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module water separator drain location" +2025-04-03 at 09:25:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:25:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank location relative to fuel cell shutdown + + +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 lunar module water separator tank location +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 water separator drain tank location before bypass line installation" +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module water separation tank location" +2025-04-03 at 09:25:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:25:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank location oxygen tank 2 vicinity + + +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 lunar module systems water separator drain tank +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "bay 4 of lunar module water separator drain location" +2025-04-03 at 09:25:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:25:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank oxygen tank 2 location + + +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: appl lunar module solid waste management system +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module bay 4 water separator drain location" +2025-04-03 at 09:26:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:26:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:26:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module water separator drain tank sensing point bay 4" +2025-04-03 at 09:26:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:26:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:26:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:26:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, False] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1636, 839, 1445, 628, 1734, 1423] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:64 - Average student length: 1284.17 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 25.00 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_correctness:66 - Length ratio: 51.37 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.688 ± 0.351 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 6.00 ± 3.11 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 4, 8, 7, 0, 9] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 09:26:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nEither a short between the temperature switch wires to ground or a contaminat...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nThe only anomaly observed in the environmental control system was a reverse l...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...'] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 09:26:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nDuring periods when the lunar module and the command module cabins were inter...', 'Result 1:\nDuring periods when the lunar module and the command module cabins were inter...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...'] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:26:02 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nIn using the lunar module water gun to dampen a towel, a piece of towel mater...', 'Result 1:\nIn using the lunar module water gun to dampen a towel, a piece of towel mater...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...'] +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:26:02 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:26:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "helium tank prelaunch rise rate space launch vehicle" +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: typical helium tank rise rates during ascent +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "space shuttle helium tank pre-launch-standby rise rate" or "space shuttle launch helium tank inflation rate +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank punch-out rise rate standard +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "prelaunch spacecraft helium tank rise rate" +2025-04-03 at 09:26:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 10 prelaunch helium tank rise rate" +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise rate helium tank launch ascent +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " space shuttle helium tank pressure rise rate comparison +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "normal helium tank rise rate range" +2025-04-03 at 09:26:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "normal helium tank pressure rise rate 7.8 psi/hour" +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate normal and abnormal +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +‌توان helium tank prelaunch-standby rise rate defined" +2025-04-03 at 09:26:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:26:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "helium tank rise rate after descent engine firing" +2025-04-03 at 09:26:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate consistency +2025-04-03 at 09:26:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "normal helium tank rise rate after first and second engine firing" +2025-04-03 at 09:26:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate normal value consistency +2025-04-03 at 09:26:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:26:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "normal helium tank rise rate after second engine firing" +2025-04-03 at 09:26:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: is helium tank rise rate normal vs non-normal +2025-04-03 at 09:26:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:26:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "normal helium tank rise rate after second engine firing standard" +2025-04-03 at 09:26:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate normal rate +2025-04-03 at 09:26:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA guidance document normal helium tank rise rate after second engine firing" +2025-04-03 at 09:26:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate normal definition +2025-04-03 at 09:26:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:26:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo Mission helium tank rise rate after engine firing max-100" +2025-04-03 at 09:26:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:26:29 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:26:29 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:26:29 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1757, 1683, 154, 401, 337, 553] +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:64 - Average student length: 814.17 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 12.00 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_correctness:66 - Length ratio: 67.85 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.567 ± 0.335 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.83 ± 3.44 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [9, 8, 2, 1, 3, 0] +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 09:26:29 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 5.0 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.833 +2025-04-03 at 09:26:29 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:26:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: manufacturing process data quality control anomalies detection +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery anisotropy +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: data anomalies in data science definition +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: signal processing anomalies math behind +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in science or technology +2025-04-03 at 09:26:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ + +2025-04-03 at 09:26:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: attitude error needle problems in spacecraft +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: exploration of manifolding effect abnormal battery performance +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission data anomalies lunar module command module +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-band antenna gyroscope sensitivity C-axis induction potentiometer +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascent oxygen tank electrical fault anomaly +2025-04-03 at 09:26:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:26:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery manifolding electrolyte bridging +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission communication problems +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Nonlinear effects in antenna system C-axis induction potentiometer X-band +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Oxygen tank 2 anomalies fan circuit electrical shorts +2025-04-03 at 09:26:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:26:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA electrolyte bridging in lithium-ion battery systems +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission communication system evaluation +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: C-axis induction potentiometer operation and specification X-band antenna requirements +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Oxygen tank 2 electrical shorts fan imaging timestamp +2025-04-03 at 09:26:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:26:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon retention screens electrolyte bridging in lithium-ion batteries +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: failed C-axis induction potentiometer malfunction induction winding ground short +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank 2 short fan circuit timestamp 55:53:20 +2025-04-03 at 09:26:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:26:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:26:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon retention screens and battery potting in electrolyte bridging +2025-04-03 at 09:26:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:26:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ground failure C-axis induction potentiometer stator primary winding short +2025-04-03 at 09:26:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrolyte bridging prevention methods in descent battery cells +2025-04-03 at 09:26:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: C-axis induction potentiometer faulty operation bias winding short +2025-04-03 at 09:26:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:26:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:26:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon retention screens and electrolyte bridging +2025-04-03 at 09:26:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:26:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:27:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrolyte bridging prevention methods using epoxy paint in battery design +2025-04-03 at 09:27:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:27:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:27:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:27:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:62 - Student lengths: [766, 1484, 944, 1953, 308, 844] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:64 - Average student length: 1049.83 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 22.00 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_correctness:66 - Length ratio: 47.72 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.529 ± 0.300 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.67 ± 3.50 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 3/6 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 10, 4, 10, 5, 0] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAlthough the standard format was followed during the deactivation and postrec...', 'Result 1:\nThe error counters for the coupling display units are used by the digital aut...'] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - ���� Searched Chunk 4: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\na. Electrolyte can leak past the Teflon retention screens installe in each ce...', 'Result 1:\na. Electrolyte can leak past the Teflon retention screens installe in each ce...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\na. Electrolyte can leak past the Teflon retention screens installe in each ce...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...'] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nAt 2-l/2 hours prior to entry, the command module was fully powered up and lu...', 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...', 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...'] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nFigure 14-6.- Recorded signal strengths during high-gain antenna operation.\n\n...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...'] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nEither a short between the temperature switch wires to ground or a contaminat...', 'Result 1:\nEither a short between the temperature switch wires to ground or a contaminat...', 'Result 1:\nElectrical shorts in the fan circuit ignited the wire insulation, causing pre...', 'Result 1:\nElectrical shorts in the fan circuit ignited the wire insulation, causing pre...', 'Result 1:\nElectrical shorts in the fan circuit ignited the wire insulation, causing pre...'] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:27:01 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:27:01 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:27:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does Z-axis accelerometer bias updating work in a complement to GPS/IMU in an aircraft navigation system? +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias update space mission +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update after update +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias update value +2025-04-03 at 09:27:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias Sample 4 +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias update time +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: REACTION accelerometer bias X-axis +2025-04-03 at 09:27:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:27:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: REACTION accelerometer bias Z-axis after update +2025-04-03 at 09:27:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:27:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: REACTION Z-axis calibration value during flight +2025-04-03 at 09:27:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:27:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: REACTION accelerometer Z-axis bias +2025-04-03 at 09:27:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:27:19 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:27:19 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:27:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:62 - Student lengths: [460, 299, 670, 610, 390, 298] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [44, 44, 44, 44, 44, 44] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:64 - Average student length: 454.50 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 44.00 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_correctness:66 - Length ratio: 10.33 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.288 ± 0.221 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 2.34 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 2, 2, 7, 0, 0] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 09:27:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:27:19 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:27:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission fuel cell performance +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space fuel cell average current Apollo missions +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Space Shuttle Discovery average fuel cell current +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the average current of the fuel cells used in a specific space mission? +2025-04-03 at 09:27:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:27:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average current fuel cell power system Apollo +2025-04-03 at 09:27:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:27:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average current of fuel cells during Apollo mission +2025-04-03 at 09:27:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo fuel cell power output +2025-04-03 at 09:27:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:27:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll 15 fuel cell average current +2025-04-03 at 09:27:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:27:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 fuel cell power consumption +2025-04-03 at 09:27:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll mission fuel cell current loss due to oxygen consumption +2025-04-03 at 09:27:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:27:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 fuel cell power consumption calculation +2025-04-03 at 09:27:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic oxygen tank failure fuel cell current loss +2025-04-03 at 09:27:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 09:27:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 fuel cell total energy supplied +2025-04-03 at 09:27:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll mission fuel cell 2 performance after oxygen tank 2 pressure loss +2025-04-03 at 09:27:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:27:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate fuel cell power consumption apollo 13 +2025-04-03 at 09:27:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll mission fuel cell 2 power duration after oxygen tank 2 pressure loss +2025-04-03 at 09:27:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:27:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average duration of a typical apollo mission +2025-04-03 at 09:27:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:27:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:27:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:27:45 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, True] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:62 - Student lengths: [424, 330, 416, 2052, 338, 361] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:64 - Average student length: 653.50 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_correctness:66 - Length ratio: 65.35 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.417 ± 0.398 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.00 ± 3.27 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 2, 0, 8, 1, 7] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:27:45 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:27:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: system temperature cold soak duration 7 hours +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What temperature did the NASA Artemis lunar lander experience during a 7-hour cold soak test? +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard temperature for cold soaking in electronic components, specifically 7 hours? +2025-04-03 at 09:27:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 09:27:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:27:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the 7-hour cold soak temperature for the Lunar Orbital Environment Experiment (LOE) apparatus on the Apollo spacecraft? +2025-04-03 at 09:27:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:27:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Find information on cold temperatures during the Apollo mission, specifically a mention of 7 hours of cold soaking, or conditions around that time frame for critical systems. +2025-04-03 at 09:27:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:27:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the highest temperature reached by the heat flow experiment during the cold soak test for the Apollo 13 experiment? +2025-04-03 at 09:27:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:27:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the duration of the cold soaking period for a specific Apollo mission or system, near 7 hours, or the temperature it was exposed to. +2025-04-03 at 09:27:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:27:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:27:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA-S-70-5864: Temperature of the Heat Flow Experiment during the cold soak test +2025-04-03 at 09:27:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:27:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the duration of a cold soak or storage temperature for electronics or critical systems, typically lasting around 7 hours. +2025-04-03 at 09:27:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 09:27:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:28:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature of the lunar surface during the cold soak test +2025-04-03 at 09:28:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:28:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standardized temperature for a cold soak, typically below -18°C, for electronic components or systems? +2025-04-03 at 09:28:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:28:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:28:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Heat flow experiment thermocouple measurements during cold soak test Apollo 13 +2025-04-03 at 09:28:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:28:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the typical temperatures during a cold soak, typically far below $40^{\circ}$ F, for electronic components or systems? +2025-04-03 at 09:28:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:28:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:28:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:28:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:28:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:62 - Student lengths: [257, 1541, 323, 700, 1218, 356] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:64 - Average student length: 732.50 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 5.00 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_correctness:66 - Length ratio: 146.50 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.375 ± 0.425 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.17 ± 2.73 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 1, 6, 6, 0] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...'] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...'] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...'] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 09:28:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:28:06 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:28:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:28:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Mars Perseverance Rover guided entry and acceleration peak +2025-04-03 at 09:28:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:28:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:28:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 guided entry acceleration peak +2025-04-03 at 09:28:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:28:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:28:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:28:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:28:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, True] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:62 - Student lengths: [423, 660, 566, 224, 876, 447] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:64 - Average student length: 532.67 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_correctness:66 - Length ratio: 33.29 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.083 ± 0.186 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.33 ± 0.75 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 0, 0, 0, 0, 2] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:28:14 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:28:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s-ivb ignition time +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission timeline S-II staging S-IVB ignition +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s ii and s ivb ignition time +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Space Shuttle S-II staging and S-IVB ignition +2025-04-03 at 09:28:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar impact time and s-ivb ignition +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB ignition and S-II staging time +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Nasa Apollo 11 S ii staging and s ivb ignition time +2025-04-03 at 09:28:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s-ivb ignition time calculation +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module mission timeline ignition S-IVB +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s ii staging and s ivb ignition time +2025-04-03 at 09:28:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:28:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 s-ivb ignition time +2025-04-03 at 09:28:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:28:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long is time between s ic and s ii ignition +2025-04-03 at 09:28:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:28:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:28:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s ivb ignition time after s ii ignition +2025-04-03 at 09:28:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:28:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:28:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:28:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:62 - Student lengths: [491, 301, 254, 635, 2320, 548] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:64 - Average student length: 758.17 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 20.00 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_correctness:66 - Length ratio: 37.91 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.388 ± 0.345 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.50 ± 2.06 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 0, 3, 0, 5, 2] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:28:42 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:28:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:28:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the approximate propellant used for the initial separation of the S-IVB from the SPS-10060 spacecraft in the Apollo 11 mission? +2025-04-03 at 09:28:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:28:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo S-IVB propellant load mass at separation" +2025-04-03 at 09:28:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:28:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:28:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:28:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:62 - Student lengths: [857, 245, 594, 1589, 571, 748] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:64 - Average student length: 767.33 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_correctness:66 - Length ratio: 85.26 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.117 ± 0.165 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 0.33 ± 0.47 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [1, 1, 0, 0, 0, 0] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 09:28:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...'] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +2025-04-03 at 09:28:51 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:28:51 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:28:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System SpaceX final flight evaluation year +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final flight evaluation date +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final flight evaluation year +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final flight evaluation year +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final flight evaluation year +2025-04-03 at 09:28:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System by Apollo mission number +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System 1969-1970 Final Flight Evaluation +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final flight evaluation year or December 1969 or Ascent Propulsion System final flight evaluation September 1970 +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System final evaluation date +2025-04-03 at 09:28:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:28:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoly 10 Ascent Propulsion System +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System December 1969 Ascent Propulsion System September 1970 +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System final flight evaluation Apollo 9 or Apollo 11 +2025-04-03 at 09:29:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:29:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System test flight evaluations December 1969 September 1970 +2025-04-03 at 09:29:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System final flight evaluation January 1970 distinct +2025-04-03 at 09:29:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:29:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System final evaluation inconsistent dates January September +2025-04-03 at 09:29:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:29:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 Ascent Propulsion System final flight evaluation date +2025-04-03 at 09:29:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:29:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-II/S-IVB ignition sequence and Ascent Propulsion System tests +2025-04-03 at 09:29:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:29:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Ascent Propulsion System final flight evaluation date +2025-04-03 at 09:29:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:29:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System S-II S-IVB engines +2025-04-03 at 09:29:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:29:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Ascent Propulsion System final flight confirmation +2025-04-03 at 09:29:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:29:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB third stage engine number 5 shutdown +2025-04-03 at 09:29:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:29:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Ascent Propulsion System final flight date and Ascent Propulsion System January 1970 +2025-04-03 at 09:29:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:29:17 | INFO | src.agent:run_tool_calls:115 - ���� Search Query: Apollo Inflight anomaly cryogenic oxygen supply +2025-04-03 at 09:29:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:29:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion System final flight evaluation correct date +2025-04-03 at 09:29:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:29:19 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:29:19 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:29:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:62 - Student lengths: [469, 146, 1849, 613, 473, 437] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:64 - Average student length: 664.50 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_correctness:66 - Length ratio: 166.12 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.567 ± 0.335 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.00 ± 3.65 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 1, 9, 2, 9, 0] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:29:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:29:19 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:29:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: official report title +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: title of official lithium-ion battery report +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report title +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report title related to Cybersecurity Readiness and Infrastructure Security report +2025-04-03 at 09:29:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:29:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 7-13 mission report table +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery performance report +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission report title +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA safe earth entry mission report Abu Dhabi to Space +2025-04-03 at 09:29:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +sufficient accuracy to permit a safe earth entry. +------ + +2025-04-03 at 09:29:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: table E-I apollo 7-13 mission reports +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module battery performance report +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 supplemental report +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: National Aeronautics and Space Administration Manned Spacecraft Center Houston Texas 1970 report +2025-04-03 at 09:29:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:29:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 immediate action report +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 MSC-02545 +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:118 - ��️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: MSC-02680 Pilot Training Evaluation Report +2025-04-03 at 09:29:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:29:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training report +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: MSC-02545 summary +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Post-Flight Operations Report +2025-04-03 at 09:29:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 in-flight operations report +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 mission report +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Post-Flight Operations Report Supplement +2025-04-03 at 09:29:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 operations evaluation report +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: MSC-02545 Apollo 13 +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Manned Spacecraft Center Supplement Report +2025-04-03 at 09:29:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:29:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Supplement Report MSC-02545 +2025-04-03 at 09:29:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:29:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Change Sheet MSC-02680 +2025-04-03 at 09:29:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:29:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:29:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:29:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, False] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:62 - Student lengths: [766, 177, 497, 478, 1722, 239] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:64 - Average student length: 646.50 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 24.00 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_correctness:66 - Length ratio: 26.94 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.504 ± 0.393 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.83 ± 4.30 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 7, 0, 7, 12, 0] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...'] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +sufficient accuracy to permit a safe earth entry. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +2025-04-03 at 09:29:52 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:29:52 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:29:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel transposition docking extraction +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel space missions transposition docking extraction +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What fuel types were used in the Space Shuttle's transposition, docking, and extraction maneuvers? +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel usage space exploration transposition docking extraction +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how much reaction control fuel was used for transposition, docking, and extraction in Apollo missions +2025-04-03 at 09:29:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:29:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA attitude control fuel requirements for docking maneuver +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: total reaction control fuel usage Apollo transposition docking extraction +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the specific fuel used for reaction control in the Space Shuttle's transposition, docking, and extraction maneuvers? +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel usage orbital maneuver +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel usage in Apollo command and lunar modules +2025-04-03 at 09:30:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel consumption attitude control lunar module docking +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reaction control fuel usage for transposition, docking, and extraction in the Space Shuttle +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control attitude control transposition docking extraction +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module reaction control fuel usage +2025-04-03 at 09:30:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module reaction control system fuel +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel used for attitude control docking +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: total reaction control fuel usage lunar module Apollo mission +2025-04-03 at 09:30:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:30:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Reaction Control System fuel consumption +2025-04-03 at 09:30:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:30:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel attitude control docking +2025-04-03 at 09:30:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control fuel docking +2025-04-03 at 09:30:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:30:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system A propellant usage during docking +2025-04-03 at 09:30:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:30:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction model A propellant usage extraction +2025-04-03 at 09:30:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:30:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:30:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system A propellant incident +2025-04-03 at 09:30:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:30:19 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:30:19 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:30:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, False, True, True] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:62 - Student lengths: [673, 424, 741, 1384, 28, 636] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:64 - Average student length: 647.67 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_correctness:66 - Length ratio: 71.96 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.642 ± 0.328 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.83 ± 2.79 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 2, 3, 9, 0, 4] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Fuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zer， lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:30:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nFuel, 1b Oxi di zer， lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...'] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 09:30:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...'] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:30:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nA descent propulsion system maneuver to reestablish a free-return trajectory ...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...'] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loaded（ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 09:30:19 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:30:19 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:30:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"heart rate astronaut lunar module pilot apollo mission" + +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'Apollo 11 lunar module pilot heart rate' +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 nasa medical records heart rate +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Lunar Module Pilot heart rate during Apollo mission" +2025-04-03 at 09:30:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"heart rate lunar module pilot lunar module apollo mission" + +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'Lunar Module Pilot heart rate during Apollo 11' +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar module pilot heart rate 125 beats/min +2025-04-03 at 09:30:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:30:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'Lunar Module Pilot heart rate range Apollo 11' +2025-04-03 at 09:30:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 crew heart rate 125 beats/min armstrong +2025-04-03 at 09:30:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'Lunar Module Pilot heart rate range in Apollo 11' +2025-04-03 at 09:30:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 average heart rate neil armstrong +2025-04-03 at 09:30:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: '(Lunar Module Pilot's heart rate range during Apollo 11) or (Lunar Module Pilot's heart rate range Apollo 11 mission)' +2025-04-03 at 09:30:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: neil armstrong baseline heart rate +2025-04-03 at 09:30:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'average Lunar Module Pilot heart rate Apollo 11' +2025-04-03 at 09:30:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: neil armstrong average heart rate +2025-04-03 at 09:30:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'calculate average heart rate lunar module pilot apollo 11' +2025-04-03 at 09:30:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average heart rate for adult male +2025-04-03 at 09:30:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:30:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'heart rate formula for astronaut' +2025-04-03 at 09:30:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:30:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 'normal values for heart rate in astronauts' +2025-04-03 at 09:30:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:30:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:30:45 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, False] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:62 - Student lengths: [296, 1660, 630, 315, 535, 247] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:64 - Average student length: 613.83 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 2.00 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_correctness:66 - Length ratio: 306.92 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.458 ± 0.391 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.17 ± 3.53 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 9, 0, 0, 7, 1] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:30:45 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:30:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo 11 translunar injection angle' +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How did the Apollo 11 spacecraft achieve translunar injection angle?: Apollo 11 Translunar Injection Maneuver (TNIM) +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the space-fixed flight-path angle at translunar injection during the Apollo 11 mission? +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mission translunar injection flight-path angle +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What historical data is there on the translunar injection angles for a specific spacecraft like Apollo 11 or Gemini missions? +2025-04-03 at 09:30:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:30:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo translunar SIVB injection sun-aligned orbit' +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection flight-path angle at 72 hours +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the initial translunar injection flight-path angle of the Apollo 11 spacecraft? +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mission translunar injection angle checks maneuver feasibility +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical translunar injection azimuth for a typical uncrewed Apollo mission? +2025-04-03 at 09:30:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:30:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'apollo mission translunar injection pitch angle' +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection angle after first transearth midcourse correction +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 zero launch angle to translunar injection angle +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB translunar injection theta angle +2025-04-03 at 09:30:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:30:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo S-IVB translunar prograde attitude control angle' +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 spacecraft translunar injection angle before correction +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 initial translunar injection alignment accuracy +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SIVA S-IVB translunar injection velocity and attitude theta angle +2025-04-03 at 09:31:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'apollo translunar injection SIVA prograde angle' +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection angle before platform alignment +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 initial translunar injection angle zero tolerance +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo SIVB translunar injection theta angle at injection +2025-04-03 at 09:31:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar module pitch angle translunar injection' +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection angle before alignment check +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 space-fixed translunar injection throttle 10 percent +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo SIVB S-IVB zero theta angle +2025-04-03 at 09:31:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:31:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo translunar S-IVB injection attitude angle' +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 spacecraft translunar injection angle before separation +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection thrust level +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB zero theta angle at injection 18-20 degrees +2025-04-03 at 09:31:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:31:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar transfer orbit angle translunar injection' +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection angle before transearth phase +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection spacecraft power status +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo SIVB translunar injection angle 18-20 degrees +2025-04-03 at 09:31:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo S-IVB translunar injection angle' +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB orbit altitude and attitude during translunar injection +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 power consumption during translunar injection +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo SIVB platform alignment sun vectors +2025-04-03 at 09:31:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:31:16 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:31:16 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:31:16 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, True, True, True] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1894, 1894, 1983, 632, 897, 1022] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:64 - Average student length: 1387.00 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_correctness:66 - Length ratio: 231.17 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.638 ± 0.352 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 8.17 ± 6.41 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [20, 9, 9, 9, 0, 2] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...'] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nAs on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return t...'] +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:31:16 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:31:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what alignment is observed with Earth at time 10:43am- +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: astronomy alignment at 40.43 degrees +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the function of a star tracker in digital video processing? +2025-04-03 at 09:31:22 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Abh orbital maneuver time 40:43 inertial measurements +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: INavigation reference star may not be relevant but reference guidance star 40.43 +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the "star used" in the coarse alignment of the lunar module platform at 40:43? +2025-04-03 at 09:31:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: trimmed inertial navigation alignment 133-l/2 hours +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: кораб guide star alignment celestial reference +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the basis for determining the alignment to the sun in spacecraft? +2025-04-03 at 09:31:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:31:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: coarse alignment time 105 hours +2025-04-03 at 09:31:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the use of moon as reference in star tracker/ alignment in space mission? +2025-04-03 at 09:31:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:31:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: denebola star, 40:43 +2025-04-03 at 09:31:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:31:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the moon's visibility impact navigation and alignment in space missions? +2025-04-03 at 09:31:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:31:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection second midcourse correction 40:43 +2025-04-03 at 09:31:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:31:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical method used to determine alignment accuracy in space missions? +2025-04-03 at 09:31:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:31:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the role of the inertial measurement unit in spacecraft attitude determination? +2025-04-03 at 09:31:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:31:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the limitations of the inertial measurement unit in spacecraft navigation? +2025-04-03 at 09:31:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:31:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:31:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the impact of IMU misalignment on spacecraft navigation and attitude control? +2025-04-03 at 09:31:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 09:31:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:31:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:31:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, False] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:62 - Student lengths: [553, 2056, 298, 1615, 695, 43] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:64 - Average student length: 876.67 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 15.00 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_correctness:66 - Length ratio: 58.44 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.346 ± 0.373 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.17 ± 3.62 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 7, 3, 9, 0, 0] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 09:31:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe entry attitude and platform alignment were confirmed by a successful sext...', 'Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...', 'Result 1:\nEarth-centered inertial coordinates. Bystem. **Change in velocity showm in bo...', 'Result 1:\nInitial outside observations through the lunar module windows indicated that ...', 'Result 1:\nInitial outside observations through the lunar module windows indicated that ...', 'Result 1:\nThe error counters for the coupling display units are used by the digital aut...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nTo conserve reaction control fuel when holding an attitude, a wide deadband w...'] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug，36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +2025-04-03 at 09:31:44 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:31:44 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:31:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:31:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what are sites 8 and 9 with recorders +2025-04-03 at 09:31:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:31:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why recorders were started hours before launch date +2025-04-03 at 09:31:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:31:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:31:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why were recorders at sites 8 and 9 started several hours before Apollo 13 launch +2025-04-03 at 09:31:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:31:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reason Apollo 13 recorders sites 8 and 9 started hours before launch +2025-04-03 at 09:31:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:31:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:31:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why were recorders started at sites 8 and 9 several hours before Apollo 13 launch +2025-04-03 at 09:31:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:31:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: purpose of recorders at sites 8 and 9 +2025-04-03 at 09:32:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what were the recorders for sites 8 and 9 installed for +2025-04-03 at 09:32:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: purpose of Apollo S-IVC recorders +2025-04-03 at 09:32:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:32:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB monitoring recorders +2025-04-03 at 09:32:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:32:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:32:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:32:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:62 - Student lengths: [558, 1869, 427, 582, 828, 1379] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [36, 36, 36, 36, 36, 36] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:64 - Average student length: 940.50 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 36.00 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_correctness:66 - Length ratio: 26.12 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.163 ± 0.230 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 1.83 ± 3.29 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 9, 2, 0, 0, 0] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +2025-04-03 at 09:32:06 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 2.0 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.333 +2025-04-03 at 09:32:06 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:32:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rate of change (decrease) in manifold pressure phenomenon areas like aerospace engineering +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the duration of decrease in system-A propellant manifold pressures during the Artemis I mission? +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Raptor engine HPA and LPA propellant manifold pressure decrease duration +2025-04-03 at 09:32:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II engine ignition propellant manifold pressure recovery +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: participant duration of decrease 22 psi system-A propellant manifold pressures Artemis I +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: system-A propellant manifold pressure decrease duration +2025-04-03 at 09:32:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:32:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propellant manifold pressure decrease duration +2025-04-03 at 09:32:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: definition of 4-5 second system-A propellant manifold pressure decrease duration +2025-04-03 at 09:32:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:32:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:32:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system activation duration of propellant pressure decrease +2025-04-03 at 09:32:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:32:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system firing duration midcourse correction +2025-04-03 at 09:32:25 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:32:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: midcourse correction duration +2025-04-03 at 09:32:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-03 at 09:32:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:32:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:32:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, False] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1801, 545, 454, 1376, 224, 643] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:64 - Average student length: 840.50 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 14.00 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_correctness:66 - Length ratio: 60.04 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.242 ± 0.259 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.33 ± 2.87 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 0, 3, 0, 3, 0] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 09:32:27 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:32:27 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:32:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: status of Apollo 10 photography report +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 mission visual inspection results +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'status Apollo 10 photography report' +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site:www.nasa.gov "Apollo 10 Mission Report" + status + update +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 photography report analysis visual inspection outcome +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: details of apollo 10 photography report analysis +2025-04-03 at 09:32:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: status of Apollo 10 Photography and Visual Information 1969-1970 +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 visual inspection results Apollo 11 +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'status Apollo 10 Photography and Visual Observations In report' +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site:nasa.gov ( Apollo 10) "status" + "analysis of ascent and descent propulsion systems" +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 final report revision status +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: status of Analysis of Apollo 10 Photography and Visual Observations In +2025-04-03 at 09:32:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reason for cancellation of Apollo 10 photography report +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 NASA documentation +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'final evaluation Apollo 10 Photography and Visual Observations In report' +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site:nasa.gov "Apollo 10 Photography and Visual Observations report" + correlation + March 1970 +2025-04-03 at 09:32:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 mission status +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many copies of Apollo 10 supplement numbers 9 are present +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ' publication date Apollo 10 Photography and Visual Observations In' +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site:nasa.gov "Apollo 10" + final +2025-04-03 at 09:32:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:32:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the official title of Apollo 10 analysis +2025-04-03 at 09:32:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'initial report Apollo 10 Photography and Visual Observations In' +2025-04-03 at 09:32:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:32:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'first report Apollo 10 Photography and Visual Observations In' +2025-04-03 at 09:32:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo Mission Study Apollo 10 report' +2025-04-03 at 09:32:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:32:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'Apollo 10 Mission Report' +2025-04-03 at 09:32:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:32:55 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:32:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, True] +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:62 - Student lengths: [875, 927, 2038, 745, 851, 286] +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:64 - Average student length: 953.67 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_correctness:66 - Length ratio: 105.96 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.600 ± 0.202 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.50 ± 2.29 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 3/6 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [4, 7, 8, 8, 2, 4] +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 6.0 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 1.000 +2025-04-03 at 09:32:55 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:32:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the time the O.J. simpson police pursuit with the white bronco ended +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What time a compressor in a refrigerator or a vacuum system is typically turned off? +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: time suit compressor event +2025-04-03 at 09:33:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:33:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the name of the Apollo mission that experienced the accident described +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 metric suit compressor turnoff time +2025-04-03 at 09:33:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:33:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the time the oxygen tank rupture occurred +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 suit compressor turnoff time Wiki +2025-04-03 at 09:33:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:33:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 oxygen tank rupture time +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 suit compressor shutdown time +2025-04-03 at 09:33:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:33:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the name of the document related to the oxygen tank anomaly report +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oxygen tank failure" and "lunar module suit compressor turned off time" +2025-04-03 at 09:33:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:33:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the approximate time of the oxygen tank 2 pressure drop +2025-04-03 at 09:33:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:33:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:33:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: was the suit compressor turned off during the described oxygen tank 2 anomaly +2025-04-03 at 09:33:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:33:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:33:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query +2025-04-03 at 09:33:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:33:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:33:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:33:21 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:62 - Student lengths: [591, 149, 765, 591, 545, 472] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:64 - Average student length: 518.83 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_correctness:66 - Length ratio: 64.85 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.454 ± 0.391 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 3.67 ± 3.50 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 1, 0, 0, 8, 5] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +(section ll.3). +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 09:33:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 La...'] +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:33:21 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:33:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: earth return mission AS-102 +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission 102 launch location +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "AS-102 mission location" +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: task shuttle automated plugs mission location +2025-04-03 at 09:33:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:33:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Gemini 12 mission +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 102 launch location +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo-AS-102 mission" +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manned Spacecraft Center Houston September 1970 AS-102 mission location +2025-04-03 at 09:33:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:33:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Gemini 12 aborted lunar orbit +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 command module recovery +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 recovery operations AS-102" +2025-04-03 at 09:33:33 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Gemini 12 astronauts +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ + +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 hoa command separation +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 recovery AS-102 mission" +2025-04-03 at 09:33:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:33:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 separation from cr's +2025-04-03 at 09:33:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 and 13 AS-508 AS-102 instrument /system" +2025-04-03 at 09:33:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:33:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 recovery site +2025-04-03 at 09:33:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 7 8 9 12 AS-508 AS-102" +2025-04-03 at 09:33:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:33:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: usso iwo jima recovery apollo 12 +2025-04-03 at 09:33:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:33:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 12 AS-508 AS-102" +2025-04-03 at 09:33:42 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:33:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: a uss iwo jima recovery mission apollo 12 +2025-04-03 at 09:33:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 AS-508 AS-102" +2025-04-03 at 09:33:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:33:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: uss iwo jima recovery locations in samoa or new zealand +2025-04-03 at 09:33:45 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:33:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:33:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:33:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.17 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:62 - Student lengths: [578, 1642, 1957, 510, 499, 472] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:64 - Average student length: 943.00 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 18.00 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_correctness:66 - Length ratio: 52.39 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.425 ± 0.363 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.17 ± 3.72 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [4, 9, 9, 3, 0, 0] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ + +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nMi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-...', 'Result 1:\nMi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...', 'Result 1:\nThe flight crew remained aboard the Iwo Jima overnight and were flown to Pago...'] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ["Result 1:\nThe Iwo Jima's position was established accurately using a satellite navigati...", 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +2025-04-03 at 09:33:46 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:33:46 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:33:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module Hickam Air Force Base to Downey, California duration +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long did it take for Apollo 8 to fly from Hickam AFB, Hawaii to California, covering the distance from Hickam to the manufacturer's plant at Downey? +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long did Apollo 11 command module travel from Hickam Air Force Base to Downey, California +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Apollo 12 command module Hickam to Downey California travel time +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 command module flight from Edwards Air Force Base to Downey +2025-04-03 at 09:33:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module Iwo Jima to Downey California flight duration +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the expected distance between Hickam Air Force Base in Hawaii to Downey, California, and how long did it take for a flight to cover this distance in a real-life mission? +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long was flight to Downey, California for Apollo 11 command module after launch from Hickam Air Force Base +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Hickam to Downey California command module travel duration based on test data +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 launch date Hickam Air Force Base +2025-04-03 at 09:33:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:33:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module flight to Downey length of flight +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 20, 1970 - april 27, 1970 duration of flight from Hickam air force base +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Hickam to Downey California command module first midcourse correction timing +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 launch date +2025-04-03 at 09:34:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module flight duration HI to Downey +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flight duration from hickam air force base to downey california using c-133 aircraft +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth trip duration estimates for Apollo 12 +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo 13 circumlunar flight duration +2025-04-03 at 09:34:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:34:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module duration flight Hickam to Downey +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 25, 1970 flight duration +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth trajectory analysis for Apollo 12 to confirm flight duration Hickam to Downey +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: speed of command module during circumlunar phase +2025-04-03 at 09:34:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:34:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: june 5, 1970 approximate arrival in downey california +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth flight time Apollo 12 sequence to confirm duration +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch time Hickam Air Force Base +2025-04-03 at 09:34:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:34:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Aside from the latitude and longitude await calculation of distance +2025-04-03 at 09:34:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 09:34:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth injection time to partial lunar orbit confirm duration +2025-04-03 at 09:34:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:34:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:34:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: haversine formula command module distance from hickam air force base +2025-04-03 at 09:34:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth injection to landing time Hickam to Downey California confirm duration +2025-04-03 at 09:34:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:34:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:34:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: downey california from hickam air force base 2.5 days +2025-04-03 at 09:34:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:34:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Transearth flight duration and crew time Hickam to recovery confirm +2025-04-03 at 09:34:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:34:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 26, 1970 departure from hickam air force base +2025-04-03 at 09:34:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:34:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:34:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:34:21 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, True, False, False, False] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:62 - Student lengths: [795, 350, 604, 1543, 1182, 1984] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:64 - Average student length: 1076.33 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_correctness:66 - Length ratio: 67.27 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.558 ± 0.326 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.83 ± 3.85 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 2, 0, 10, 10, 8] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe transearth injection maneuver was performed on time, and the transearth c...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:21 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nManeuver System Ignition time, hr:min:sec Firing time; sec Velocity change. f...', 'Result 1:\nEquipment transfer and stowage in both the command module and lunar module wa...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:34:21 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:34:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the name of the mission or spacecraft that is known for landing on a specific celestial body? + + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Artemis mission landing site +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Curiosity Rover landing site +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the location of the first lunar landing or a notable space mission landing? +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: arsenals spacecraft landing location +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Perseverance rover landing site Jezero crater +2025-04-03 at 09:34:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:34:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manned Spacecraft Center Houston 1970 +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of Apollo 8 splashdown +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: task Apollo 11 landing site +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing site coordinates +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: uncrewed phased radar 422 recovery operations Iwo Jima +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module descent engine firing +2025-04-03 at 09:34:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing site +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 25.5 degrees south latitude 176 degrees west longitude +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission post-flight analysis +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module landing site +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: recovery operations Uncrewed Phased Radar (UPR) Hawaii +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power system issues Apollo 11 +2025-04-03 at 09:34:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:34:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar landing site +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 21.38°N 165.2432°W +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Not chosen, What about Apollo 11 mission documentation +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Uncrewed Phased Radar (UPR) landing latitude and longitude +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module powered-down state +2025-04-03 at 09:34:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:34:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar landing coordinates +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 21° N 165° W +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission documentation +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module jettison location +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Uncrewed Phased Radar (UPR) mission landing coordinates +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 landing site altitude +2025-04-03 at 09:34:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 landing site +2025-04-03 at 09:34:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 landing site Fra Mauro Highlands +2025-04-03 at 09:34:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:34:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:34:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 landing site Apollo 12 +2025-04-03 at 09:34:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:34:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission description +2025-04-03 at 09:34:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:34:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing +2025-04-03 at 09:34:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:34:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 attempted landing site +2025-04-03 at 09:34:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:53 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:34:53 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:34:53 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, True, True, False] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:62 - Student lengths: [258, 317, 1956, 382, 238, 1956] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:64 - Average student length: 851.17 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 17.00 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_correctness:66 - Length ratio: 50.07 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.792 ± 0.224 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 6.33 ± 1.97 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 5, 10, 5, 5, 8] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe lunar module was jettisoned l hour before entry, which was performed nomi...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', "Result 1:\nThe Iwo Jima's position was established accurately using a satellite navigati...", 'Result 1:\n+21.0 +3.0\n------\nResult 2:\nTABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazin...', 'Result 1:\n+21.0 +3.0\n------\nResult 2:\nTABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazin...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe lunar module was jettisoned l hour before entry, which was performed nomi...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nTABLE 1O.3-I.- RECOVERY SUPPORT\n\nLanding area Supporta Remarks Number Unit La...', 'Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...', 'Result 1:\nThe first reported electronic contact by the recovery forces was through S-ba...', 'Result 1:\nThe updata link was used when required and performed nominally. No VHF equipm...', 'Result 1:\nThe updata link was used when required and performed nominally. No VHF equipm...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:34:53 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\na. Perform selenological inspection, survey, and sampling of materials in a p...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:34:53 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:34:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: search for the material used in the bristles of a vacuum brush and its potential replacement material. +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what vacuum brush bristles were replaced by > +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were the bristles on some vacuums changed from? + + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were Dyson vacuum bristles replaced from? +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What modern vacuum model changed bristles to softer materials on its vacuum brush? +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were the bristles on the GE Energy Star vacuum brush changed from? +2025-04-03 at 09:34:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:34:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: information about the material used for vacuum brush bristles that were changed from Teflon to Nylon +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: _before change to nylon > vacuum brush bristle material +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were vacuum brush bristles originally made of before they were changed for space suits? + + +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vacuum cleaner bristle material used in modern robot vacuum models +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Based on the context of vacuum brushes in general) What space vacuum cleaner brushes used to be made from Teflon before changing to Nylon? +2025-04-03 at 09:35:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum brush bristles material used by NASA +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of bristles were originally used on vacuum brushes before being replaced with nylon in the Apollo program? + + +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA lunar brushing brush vacuum cleaner model +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the original material of vacuum brushes used on early Apollo missions? +2025-04-03 at 09:35:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:35:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: origin of vacuum brush bristles +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum brush bristle material change to +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were bristle vacuum brush components originally made of? + + +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA lunar dust reducing vacuum brush vacuum cleaner model +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Based on the context of vacuum brushes in general) What type of fiber was used to replace Teflon in vacuum brushes for vacuum cleaning applications? +2025-04-03 at 09:35:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stanley vacuum brush bristle material +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum brush bristle material Teflon to Nylon +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were bristle vacuum cleaner bristles originally made of before being changed to nylon? + + +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA space suit vacuum brush material switched to nylon +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of synthetic fiber was used to replace Teflon in household vacuum cleaner brushes? +2025-04-03 at 09:35:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stanley vacuum brush bristle material +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon nylon brush contamination remediation +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material are traditional vacuum brush bristles typically made of? + + +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space suit vacuum brush nylon bristles +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What synthetic fiber replaced the original filament material in vacuum cleaner bristle brush used for small appliances? +2025-04-03 at 09:35:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stanley vacuum brush bristle material specifications +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon nylon vacuum brush hose valve issue +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of fabric are traditional vacuum cleaner brush bristles typically made of? + + +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space suit vacuum brush model +2025-04-03 at 09:35:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How is the vacuum brush or vacuum brush bristles used in a Stanley vacuum sealer +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon nylon vacuum vent valve +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material were vacuum brush bristles changed from for space suits? + + +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA changed vacuum brush Bristles to Nylon for contamination +2025-04-03 at 09:35:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stanley vacuum sealer brush replacement +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nylon vacuum system postlanding +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the original material of vacuum brush bristles? + + +2025-04-03 at 09:35:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stanley vacuum sealer brush material +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nylon vacuum brush postlanding +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the material vacuum cleaner bristles were originally made of? + + +2025-04-03 at 09:35:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:35:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:35:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, False, False, True, False] +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1125, 755, 689, 383, 326, 72] +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:64 - Average student length: 558.33 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 6.00 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_correctness:66 - Length ratio: 93.06 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.850 ± 0.224 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 7.50 ± 3.25 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [10, 10, 10, 1, 8, 6] +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 6.0 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 1.000 +2025-04-03 at 09:35:18 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:35:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module operations kennedy space center history display +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center lunar module history exhibit location +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission control center ksc archives +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: LM-7 Apollo 7 mission documentation Kennedy Space Center +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module LM-7 operations history Kennedy Space Center Apollo +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo-Saturn V exhibit Kennedy Space Center +2025-04-03 at 09:35:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:35:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (no relevant data found) apollo program history lunar module operations +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center Apollo 9-12 lunar module operations history +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 and Apollo 9 LM-7 documentation +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module operations Kennedy Space Center +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: abort profile Apollo 13 lunar module +2025-04-03 at 09:35:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:35:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module operations +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module operations history +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Lunar Module Abort Guidance +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module recovery Kennedy Space Center operations +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo LM-7 Apollo 13 +2025-04-03 at 09:35:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:35:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lm 7 training +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing training location +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Abort Guidance System procedures +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 LM-6 observations Kennedy Space Center +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew lunar module training +2025-04-03 at 09:35:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 7 lunar module history +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center Apollo 13 training and operations +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Descent Propulsion System Transearth Injection Apollo 9 +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Lunar Module operations Kennedy Space Center +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo LM-7 display at Kennedy Space Center +2025-04-03 at 09:35:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lm7 lunar module +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center Apollo 13 mission support and lunar module operations +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Lunar Module operations at Kennedy Space Center +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Located history of Apollo LM-7 at Kennedy Space Center +2025-04-03 at 09:35:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:35:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: history of lunar module programs +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manuals or documents covering the Apollo 13 mission operations at Kennedy Space Center for lunar module +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center Apollo 11 command module operations +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module display history +2025-04-03 at 09:35:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:35:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module operations Kennedy Space Center +2025-04-03 at 09:35:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center Apollo missions exhibits +2025-04-03 at 09:35:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:35:55 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:35:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, False, False] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.17 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:62 - Student lengths: [415, 2058, 500, 836, 2052, 2049] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:64 - Average student length: 1318.33 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_correctness:66 - Length ratio: 131.83 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.612 ± 0.243 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 6.50 ± 2.63 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 3/6 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 7, 1, 6, 9, 8] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nAfter powering up the lunar module, co-aligning the two platforms, and shutti...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAfter powering up the lunar module, co-aligning the two platforms, and shutti...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:35:55 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nBecause of a sudden loss of pressure at approximately 56 hours from one of th...', 'Result 1:\nMedical kits for future flights will include nose drops packaged the same as ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:35:55 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:35:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum rate of climb +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"pitch gimbal actuator maximum rate excursion degrees per second" + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum rate excursion +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal limits or specifications for the DJI Mini 2 electric quadcopter max climb rate in degrees per second +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum rate of turn in degrees per second +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: maximum rate excursion of DJI Mavic pitch gimbal in degrees per second +2025-04-03 at 09:36:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shuttle tempo > shuttleroll/pitch gimbal actuator max velocity +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"pitch gimbal actuator maximum excursion rate in degrees per second" + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum excursion +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: yaw rate limits operation angle limit for the Dragonfly +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gimbal pitch drive actuator angular rate specification +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum rate +2025-04-03 at 09:36:06 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"degrees per second roll gimbal drive actuator excursion" + +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator angular velocity specification +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum excursion rate +2025-04-03 at 09:36:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal angular velocity rate +2025-04-03 at 09:36:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gimbal actuator roll excursion rate +2025-04-03 at 09:36:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:36:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum rate of excursions +2025-04-03 at 09:36:15 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:36:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal drive actuator maximum excursion rate +2025-04-03 at 09:36:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:36:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch gimbal actuator rate of change +2025-04-03 at 09:36:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:36:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:36:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, False, False, True] +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:62 - Student lengths: [294, 890, 429, 640, 307, 1996] +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:64 - Average student length: 759.33 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_correctness:66 - Length ratio: 189.83 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.542 ± 0.149 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.00 ± 2.77 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 3, 2, 3, 4, 10] +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe first midcourse correction maneuver, performed at the second option point...', 'Result 1:\nThe first midcourse correction maneuver, performed at the second option point...'] +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 5.0 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.833 +2025-04-03 at 09:36:18 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:36:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the status of item number 9 recently or historically mentioned in various categories? +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the current status of the 9th item in the list? +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the status of item 9 +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the status of item 9? +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Beatles ninth studio album release date +2025-04-03 at 09:36:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 09:36:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: a table listing the status of each Apollo mission report, specifically item 9.0. +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the status of item number 9 in the Apollo 13 mission reports section "Biomedical Evaluation" unclear or not available in the provided document? +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: status of Apollo 7 through Apollo 13 mission report 9 +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical evaluation 9-1 +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: TRANSEARTH COAST 9 +2025-04-03 at 09:36:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-9 mission reports, specifically report number 9. +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Do the provided data for the biomedical evaluation and physiological parameters of the Apollo 13 crew members in Result 2 indicate that the status of item number 9, "Physical Examinations," was documented with some unknowns or issues during the mission? +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 9 flight evaluation report +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical evaluation status during apollo mission +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Trans Earth coast. +2025-04-03 at 09:36:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:36:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-9 supplemental report listing, specifically list of reports with Apollo 9 report number +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the status of the crew members' heart rates after the telemetry dropout at 55:54:54, as mentioned in Result 2, concerning weightlessness adaptation, +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cancelled apollo 10 report 9 +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical evaluation apollo 13 routine findings routine physical parameters +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth coast tm -9 +2025-04-03 at 09:36:38 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 09:36:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-13 supplemental reports, including table not in Table E-I +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the Command Module Pilot's heart rate, which decreased from 68 to 60 beats/min after the incident, and the Lunar Module Pilot's heart rate increase to 125 beats/min after the incident, change the known status of "Physical Examinations" to indicate issues or unknowns during the mission? +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 report 9 evaluation status +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module biomedical evaluation apollo 13 +2025-04-03 at 09:36:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:36:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 Supplement 14 report listing +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can the lunar module pilot's heart rate of 125 beats/min during entry indicate that the Physiological Examination was affected by the mission's severe event? +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cancelled evaluation of apollo 10 report 9 +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module helium pressure control apollo 13 +2025-04-03 at 09:36:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:36:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:36:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 report 9 cancelled status +2025-04-03 at 09:36:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module helium pressure rise rate cold soak apollo 13 +2025-04-03 at 09:36:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:36:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:36:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: official reason for cancellation of apollo 10 report 9 +2025-04-03 at 09:36:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:36:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: preflight test helium pressure rise rate apollo 13 +2025-04-03 at 09:36:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:36:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:36:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank performance apollo 13 +2025-04-03 at 09:36:50 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:36:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:36:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:36:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, True, False, False, False, True] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:62 - Student lengths: [224, 277, 235, 1958, 1777, 2054] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:64 - Average student length: 1087.50 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 9.00 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_correctness:66 - Length ratio: 120.83 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.712 ± 0.363 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 6.67 ± 3.45 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 6, 6, 8, 9, 11] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...'] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water. + +9.0 BIOMEDICAL EVALUATION + +This section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted. + +9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 09:36:50 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nGuidance system performance was again nominal and there were no significant a...', 'Result 1:\nTABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazine L frame Mission elapsed ti...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...'] +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:36:50 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:36:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: charge time of a specific battery +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery charging time rundown < +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery charge time +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: charge time for battery B +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long does it take to fully charge a typical lithium-ion battery? +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: charge time for battery B information +2025-04-03 at 09:36:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:36:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 module power system batteries charging time < +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery charging factors +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent battery capacity +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long does a fully charged battery take to reach its full capacity after a surge event? +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent-stage battery and ascent-stage battery together +2025-04-03 at 09:37:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:37:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module battery type shape firing time < +2025-04-03 at 09:37:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion battery charging time 10 minutes +2025-04-03 at 09:37:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of current surge at 97:13:56 +2025-04-03 at 09:37:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:37:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Supporting documents for battery charging and power consumption on the lunar module during Apollo mission +2025-04-03 at 09:37:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 battery 2 malfunction date time +2025-04-03 at 09:37:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module electrical battery 100-amperes surge < +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank bursting at 109 hours +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long did it take for the total energy to be delivered to the lunar module for the initial powered-down configuration after returning from the moon? +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent battery 2 malfunction time 97:13:56 +2025-04-03 at 09:37:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:37:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module battery charging time 100-amperes surge < +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module ascent tank helium pressure +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Charging time for lunar module entry batteries from start to completion for Apollo 13 mission +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery 2 event times 97:13:56 and 99:5l:09 +2025-04-03 at 09:37:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:37:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module battery 2 surge 97:13:56 charging time < +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module ascent stage pressure at 90 psia +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 97:13:56 Apr11 1970 descent descent battery 2 +2025-04-03 at 09:37:20 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lunar module surge 97:13:56 < +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure change for lunar descent +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 97:13:56 and 99:5l:09 connection descent battery 2 +2025-04-03 at 09:37:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:37:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module battery charging time_detail < +2025-04-03 at 09:37:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:26 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:37:26 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:37:26 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, True, False, True, False] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:62 - Student lengths: [387, 1938, 590, 1851, 569, 1817] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [21, 21, 21, 21, 21, 21] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:64 - Average student length: 1192.00 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 21.00 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_correctness:66 - Length ratio: 56.76 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.596 ± 0.263 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 7.17 ± 3.24 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 4/6 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [2, 8, 6, 11, 5, 11] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThree balloons containing instruments designed to measure the air/ earth curr...', 'Result 1:\nThe prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:26 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe designs of other Apollo batteries have been reevaluated, and all are cons...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...'] +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:37:26 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:37:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission timeline power down command service nodules lunar module power up +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power down and up completion time + + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar module power down and up timeline +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 command and service module power status +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the times of power down and power up for the command and service modules with lunar module Eagle after re-entry on December 21, 1969. +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 lunar module power down and power up date +2025-04-03 at 09:37:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module power-down command module powered up procedures +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Lunar module powerup was handled quite efficiently by identifying Selected segments of an existing procedure, the "Lunar Module Systems Activation Checklist." However, the crew had to delete the very high frequency portion of the communications activation. This procedure also assumed suited operations, so the crew had to turn on suit flow valves and unstow hoses to establish air flow. This extended power-up blended well with the preparation for the subsequent midcourse maneuver to enter a free return trajectory. A similar real-time update to the 2-hour activation section of the "Lunar Module Contingency Checklist" was also quite adequate. Lunar module activation was completed at the time fuel cell 2 reactant flow went to zero because of oxygen depletion. The command and service modules were then powered down completely according to a groumdgenerated procedure. To form a starting baseline for subsequent procedures, each switch and circuit breaker in the command module was positioned +------ + +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar module power down and up time 58:40:00 +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 power down and power up times +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the specific times for the power down of the command and service modules of the Apollo 11 spacecraft followed by the power up of the lunar module Eagle. +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 lunar module power down and up time +2025-04-03 at 09:37:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module power-up command module start time fuel cell 2 reactant flow to zero +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Lunar module powerup was handled quite efficiently by identifying Selected segments of an existing procedure, the "Lunar Module Systems Activation Checklist." However, the crew had to delete the very high frequency portion of the communications activation. This procedure also assumed suited operations, so the crew had to turn on suit flow valves and unstow hoses to establish air flow. This extended power-up blended well with the preparation for the subsequent midcourse maneuver to enter a free return trajectory. A similar real-time update to the 2-hour activation section of the "Lunar Module Contingency Checklist" was also quite adequate. Lunar module activation was completed at the time fuel cell 2 reactant flow went to zero because of oxygen depletion. The command and service modules were then powered down completely according to a groumdgenerated procedure. To form a starting baseline for subsequent procedures, each switch and circuit breaker in the command module was positioned +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module power down start time and command module power up start time +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What specific times were the command module powered down at 58 hours 40 minutes and 29 minutes. What was the time when the lunar module Eagle powered up after the lunar module's re-entry? +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 lunar module power up time after command module shutdown +2025-04-03 at 09:37:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: detailed timeline of power outage, undervoltage alarm, and command module recovery +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 command module power up procedure detailed +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: When was the lunar module Eagle powered up after the lunar module's re-entry on December 21, 1969. +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 descent propulsion firing time +2025-04-03 at 09:37:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 undervoltage alarm recovery time +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module second descent propulsion firing and lunar module powered down time +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 lunar module power-up time and duration after re-entry. When did the lunar module's power-up end. +2025-04-03 at 09:37:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:37:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module separation and recovery completion time +2025-04-03 at 09:37:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:37:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: When was the lunar module powered up following its re-entry on December 21, 1969. +2025-04-03 at 09:37:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:37:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:37:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 landing and recovery time +2025-04-03 at 09:37:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: When was the lunar module powered up, if it happened +2025-04-03 at 09:37:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:37:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 powered-down and powered-up lunar module and command module times +2025-04-03 at 09:37:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: entry to lunar module powered-up time +2025-04-03 at 09:37:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:37:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:37:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module powered-up time after Command Module power-down on December 21, 1969 +2025-04-03 at 09:37:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:37:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:37:56 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:62 - Student lengths: [2037, 592, 268, 286, 2037, 387] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:64 - Average student length: 934.50 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 8.00 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_correctness:66 - Length ratio: 116.81 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.696 ± 0.258 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.00 ± 2.89 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 1, 2, 5, 9, 5] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Lunar module powerup was handled quite efficiently by identifying Selected segments of an existing procedure, the "Lunar Module Systems Activation Checklist." However, the crew had to delete the very high frequency portion of the communications activation. This procedure also assumed suited operations, so the crew had to turn on suit flow valves and unstow hoses to establish air flow. This extended power-up blended well with the preparation for the subsequent midcourse maneuver to enter a free return trajectory. A similar real-time update to the 2-hour activation section of the "Lunar Module Contingency Checklist" was also quite adequate. Lunar module activation was completed at the time fuel cell 2 reactant flow went to zero because of oxygen depletion. The command and service modules were then powered down completely according to a groumdgenerated procedure. To form a starting baseline for subsequent procedures, each switch and circuit breaker in the command module was positioned +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Lunar module powerup was handled quite efficiently by identifying Selected segments of an existing procedure, the "Lunar Module Systems Activation Checklist." However, the crew had to delete the very high frequency portion of the communications activation. This procedure also assumed suited operations, so the crew had to turn on suit flow valves and unstow hoses to establish air flow. This extended power-up blended well with the preparation for the subsequent midcourse maneuver to enter a free return trajectory. A similar real-time update to the 2-hour activation section of the "Lunar Module Contingency Checklist" was also quite adequate. Lunar module activation was completed at the time fuel cell 2 reactant flow went to zero because of oxygen depletion. The command and service modules were then powered down completely according to a groumdgenerated procedure. To form a starting baseline for subsequent procedures, each switch and circuit breaker in the command module was positioned +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...'] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nEquipment transfer and stowage in both the command module and lunar module wa...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nEquipment transfer and stowage in both the command module and lunar module wa...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 4.0 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.667 +2025-04-03 at 09:37:56 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.471 +2025-04-03 at 09:37:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the ground pressure rise rate during the Countdown to Launch ( CTL) test of the Apollo program? +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 countdown demonstration test ground pressure rise rate +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nominal ground pressure rise rate during Apollo 11 countdown demonstration test +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: longest duration lithium-ion batteries for electric cars in kilotons +2025-04-03 at 09:38:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:38:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:38:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 helium tank pressure rise rate during countdown demonstration test +2025-04-03 at 09:38:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Saturn S-IVB ground test nominal pressure rise rate +2025-04-03 at 09:38:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:38:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:38:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 nominal ground pressure rise rate during countdown demonstration test +2025-04-03 at 09:38:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 helium tank pressure rise rate abnormal +2025-04-03 at 09:38:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:38:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 test with helium tank pressure rise rate 33 psi/hour +2025-04-03 at 09:38:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no change +2025-04-03 at 09:38:12 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:38:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 countdown demonstration test rise rate +2025-04-03 at 09:38:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 countdown demonstration test helium rise rate during initial cold-soak period +2025-04-03 at 09:38:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 countdown demonstration test helium pressure rise rate +2025-04-03 at 09:38:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:38:16 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:38:16 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, True, True] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:62 - Student lengths: [487, 574, 501, 1717, 712, 284] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:64 - Average student length: 712.50 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_correctness:66 - Length ratio: 71.25 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.354 ± 0.324 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 2.50 ± 3.10 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [3, 0, 0, 9, 1, 2] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:38:16 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 1.0 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.167 +2025-04-03 at 09:38:16 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.373 +2025-04-03 at 09:38:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What will secure knobs on timers on aircraft import or manufacturing +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials will be used to secure knobs on aircraft timers in future flights? +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Are aircraft timers secured with EADs or other methods? +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: secure knobs on flight timer controls aviation +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aircraft timer security standard +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: air traffic control timer security best practices +2025-04-03 at 09:38:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:38:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: roll pin vs. other securing mechanisms for aerospace applications +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of fasteners used for securing knobs on timers in aircraft +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How to secure knobs on timers or equipment handles with high safety standards +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs on timers with roll pins +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aircraft timer locking mechanism roll pin type +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: locking mechanisms for knobs +2025-04-03 at 09:38:27 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cable vs. sheet as securing material in mechanical connections +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of pins used to secure knobs on timer shafts +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securement of knobs or handles with existing materials or methods +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs with roll pins for timers in flight equipment +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-strength fasteners for securing aircraft timer knobs +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: roll pin safety valve hubs +2025-04-03 at 09:38:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mechanical valve-lock mechanism vs. electrical switch +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is a roll pin +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securing of metric used in Space Application +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs on timers alternative to roll pin +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aerospace fastener materials for securing aircraft components +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: knob security mechanisms for flight control +2025-04-03 at 09:38:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 09:38:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: secure mechanical locking mechanism for high-security applications +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what secures aircraft control panel knobs +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: secures knobs on control devices +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs on timers with electrical or mechanical connection +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aerospace gasket materials for sealants +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: secure interval timer knobs +2025-04-03 at 09:38:40 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliable mechanical linkage for high-stress or high-security applications +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: functions of roll pins +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securement of knobs on control modules +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs with electrical connectors or switches in timer systems +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal barrier materials for aerospace fasteners +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs on interval timer command module +2025-04-03 at 09:38:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliability and security considerations for high-stress applications +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: straps used on knurled knobs +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securing knobs to prevent loosening +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crimping mechanism for securing knobs +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials for high-temperature and high-pressure aerospace applications +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: securing knobs with a pin +2025-04-03 at 09:38:48 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: redundancy design for high-reliability applications +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: knurled knob fasteners +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securing knobs in high-stress applications +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crimping mechanism for securing knob terminals or switches +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gasket materials compatible with high-pressure applications Teflon +2025-04-03 at 09:38:52 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:38:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of knobs used on aircraft control panels +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securing knobs and interrupter switches +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crimped pinning or soldering for securing knobs +2025-04-03 at 09:38:55 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 09:38:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how particle knobs are secured +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Securing knobs to prevent loosening +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: soldering pins or sleeves for securing knobs +2025-04-03 at 09:38:57 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:57 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:38:57 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:38:57 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [False, False, False, True, False, True] +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.33 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1369, 1851, 1968, 1676, 1884, 1676] +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:64 - Average student length: 1737.33 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 10.00 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_correctness:66 - Length ratio: 173.73 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.871 ± 0.177 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 10.00 ± 2.00 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [8, 10, 10, 10, 8, 14] +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 6.0 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 1.000 +2025-04-03 at 09:38:57 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:38:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium-ion batteries in propulsion systems +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does an electric motor differ from a traditional internal combustion engine in terms of efficiency and operation? +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Explainer: Propulsion system types +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propulsion system types in various industries +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: specific propulsion system in aerospace engineering textbook +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of propulsion systems discussion +2025-04-03 at 09:39:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 propulsion system +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is gas turbine propulsion system +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spacecraft descent and ascent propulsion systems +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propulsion +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module earth orbit propulsion +2025-04-03 at 09:39:09 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System Apollo 13 +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system xenon gas propulsion +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Trans-Earth injection maneuver and attitude control propulsion systems +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module ascent and descent propulsion systems +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Descent Propulsion System performance +2025-04-03 at 09:39:13 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent Propulsion System in Apollo Spacecraft +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (no specific keywords are gathered) propulsion system for trans-Earth injection +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 translunar injection and S-IVB-vibration studies +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing ascent and descent systems +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Descent Propulsion type +2025-04-03 at 09:39:19 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Ascent Propulsion Systems +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (no specific keywords are gathered) propulsion system for electric or electro-mechanical spacecraft thrusters +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IC, S-II, and S-IVB engine performance issues +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar laser ranging retroreflector deployment +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Descent Propulsion System type +2025-04-03 at 09:39:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Descent Propulsion Apollo Lunar Module +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system for NASA lunar missions +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IC, S-II, S-IVB, and crewnotification of small acceleration changes during engine mixture ratio shifts +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: abort guidance system propulsion +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Descent Propulsion System performance +2025-04-03 at 09:39:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reaction Control System RCS Descent Propulsion +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Regolith Descent Engine (RDE) spacecraft propulsion +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II engine thrust and structural loads during S-II boost +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo flight attitute control system propulsion +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Descent Propulsion System solid-fueled system +2025-04-03 at 09:39:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Ascent Propulsion System +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descend only xenon propulsion systems +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IC, S-II, and S-IVB dynamic loads and structural tolerances during Apollo +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module propulsion systems comparison +2025-04-03 at 09:39:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:39:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: .Descend propulsion systems using helium gas +2025-04-03 at 09:39:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:39:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system review +2025-04-03 at 09:39:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:39:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supercritical helium pressurization for spacecraft descent propulsion systems +2025-04-03 at 09:39:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:39:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar module propulsion system evaluation +2025-04-03 at 09:39:43 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:39:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:39:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1380, 1567, 1862, 1387, 1103, 767] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:64 - Average student length: 1344.33 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 18.00 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_correctness:66 - Length ratio: 74.69 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.596 ± 0.263 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 10.50 ± 4.89 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 4/6 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [14, 10, 9, 10, 18, 2] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe first midcourse correction maneuver, performed at the second option point...', 'Result 1:\nWith the exception of supercritical helium system performance, descent propul...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe supercritical helium pressurization system displayed abnormal performance...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\nStructural loads experienced during S-IC boost were well below design values,...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...', 'Result 1:\nA descent propulsion system maneuver to reestablish a free-return trajectory ...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 09:39:43 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...'] +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:39:43 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:39:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: duplicated malfunction in spacecraft + +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reusable rockets falcon 9 unplanned emergency landing +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 oxygen tank explosion cause +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happened when malfunction duplicate error occurred in roversperussiamission Mariner spacecraft +2025-04-03 at 09:39:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 09:39:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: duplicated malfunction in Apollo spacecraft + +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what happened to falcon 9 rocket when it duplicated its malfunction +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch lightning volts charge separation +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission malfunction duplication what happened +2025-04-03 at 09:39:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:39:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: duplicated malfunctions in Apollo 12 + +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happened to the reused Falcon 9 rocket's reaction control quads after loss of control +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 electrical discharge electrostatic potential 160 meej +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 transducer malfunction duplication consequences +2025-04-03 at 09:39:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:39:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Apollo 12 malfunction end of zero optics mode + +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 spacecraft electrostatic potential 4 million volts 160000 joules +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 transducer contamination and pressure issues fix +2025-04-03 at 09:39:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:39:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: Apollo 12 re-entry guidance navigation and control system malfunction + +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 saturn v electrostatic potential comparison +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 transmission system cause +2025-04-03 at 09:39:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 09:39:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: navigation and guidance system malfunctions Apollo 13 re-entry phase + +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: saturn v electrostatic potential in comparison to conventional aircraft +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 supersonic altimeter malfunction +2025-04-03 at 09:40:01 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: helium pressure stabilization lunar module Apollo 13 + +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 saturn v much higher than million volts +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission post abort troubleshooting +2025-04-03 at 09:40:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: helium pressure regulation system testing Apollo 13 mission + +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 radio noise onset lift off +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 emergency abort procedures +2025-04-03 at 09:40:05 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: helium pressure stabilization system mitigation measures + +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 radio noise timing electric field perturbation +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ + +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 emergency abort successful recovery +2025-04-03 at 09:40:08 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 09:40:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:40:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 radio noise electric field perturbation simultaneous +2025-04-03 at 09:40:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ + +2025-04-03 at 09:40:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 orbital data +2025-04-03 at 09:40:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:40:10 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:40:10 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:40:10 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, False, False] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1657, 481, 594, 1770, 164, 1956] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [33, 33, 33, 33, 33, 33] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:64 - Average student length: 1103.67 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 33.00 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_correctness:66 - Length ratio: 33.44 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.583 ± 0.426 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.33 ± 4.46 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [9, 3, 0, 10, 0, 10] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\na. Some of the radioluminescent disks were broken. b. The apex cover was brok...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThis transducer has operated erratically on two previous missions. Testing af...', 'Result 1:\nTo guard against operational problems of this type in the future, a caution n...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe supercritical helium pressurization system displayed abnormal performance...'] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nFollowing the anomaly, all reaction control quads except C began showing evid...', 'Result 1:\nFollowing the anomaly, all reaction control quads except C began showing evid...'] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz. + +During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids. +------ + +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nTo monitor the low-frequency radio noise, a broad-band antenna system was use...', 'Result 1:\nBecause of access restrictions to sites 8 and 9, the corresponding recorders ...', 'Result 1:\nBecause of access restrictions to sites 8 and 9, the corresponding recorders ...'] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:40:10 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThis transducer has operated erratically on two previous missions. Testing af...', 'Result 1:\nThis transducer has operated erratically on two previous missions. Testing af...', 'Result 1:\nInspection also revealed that both the cabin and suit loop pressure transduce...', 'Result 1:\nThis transducer has operated erratically on two previous missions. Testing af...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...', 'Result 1:\nMedical kits for future flights will include nose drops packaged the same as ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:40:10 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:40:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What data types are commonly recorded in the last row of a table +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of data is recorded in the final entry of a database table? +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What data types are typically recorded in the last entry of a table? +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type of data in last entry of table +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of data are typically recorded in a table's last entry? +2025-04-03 at 09:40:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: data type of information in tables D-I and E-I +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: data types in table entry l +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: availability data format command module +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what type of data are reported in table E-I for mission reports +2025-04-03 at 09:40:18 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: contents of tables D-I and E-I in Apollo mission documentation +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: definition of reports in a table +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: tabular data format command module status +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Table D-I contents +2025-04-03 at 09:40:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descriptions and table headings in appendix sections of Apollo mission documentation +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: content of written mission reports +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: categorical data table format checkmark or no +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: difference in data focus between D-I and E-I +2025-04-03 at 09:40:23 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:40:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission performance analysis reports +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: written human-document +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(section ll.3). +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: binary coding system yes no table +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: iris data collected during lunar mission +2025-04-03 at 09:40:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:40:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:40:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pre Apollo 12 mission report table naming convention +2025-04-03 at 09:40:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:40:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: identification of the type of data in the experimental component list +2025-04-03 at 09:40:29 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:40:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:40:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: table format checkmark or yes/no with categorical headers +2025-04-03 at 09:40:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:40:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type of data in DATA AVAILABILITY section +2025-04-03 at 09:40:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:40:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: information on Table D-II listings for lunar module +2025-04-03 at 09:40:34 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:40:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:40:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: description of fields in A.2 LUNAR MODULE in D-II +2025-04-03 at 09:40:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:40:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when was Table D-II listed in the released documentation +2025-04-03 at 09:40:37 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 09:40:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:40:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:40:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, False, True] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.67 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.47 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:62 - Student lengths: [812, 376, 79, 385, 517, 1157] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:64 - Average student length: 554.33 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 4.00 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_correctness:66 - Length ratio: 138.58 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.692 ± 0.379 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.67 ± 3.40 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 0/6 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [5, 5, 1, 7, 10, 0] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +(section ll.3). +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\n(section ll.3).\n------\nResult 2:\ndiscussion of the flight and associated crew...'] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...'] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: ++21.0 +3.0 +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nTables D-I and D-II are summ...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nAlthough the standard format was followed during the deactivation and postrec...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\n+21.0 +3.0\n------\nResult 2:\nAlthough the standard format was followed during ...'] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +DESCRIPTIONS·········· A-1 A.1 COMMAND AND SERVICE MODULES .···· A-1 A.2 LUNAR MODULE ······· A-1 A.3 EXPERIMENT EQUIPMENT · A-2 A.4 LAUNCH VEHICLE ······ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\napproximately l mile of the target location. Because of power restrictions, t...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nTables D-I and D-II are summaries of the da...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...'] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 09:40:37 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:40:37 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:40:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: consider components and connections in a vehicle's electrical system where a short circuit could occur +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is a momentary short circuit in DC electrical system +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliable sources of information for locating short circuit faults in electrical systems +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of a momentary short circuit in a DC power distribution system, particularly in the context of electrical engineering and power systems? +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: locations of short circuits in batteries +2025-04-03 at 09:40:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:40:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of electrical short circuit interrupters on spacecraft electrical system +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent stage dc system connection +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: short circuit location C-axis induction potentiometer reaction control system +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent battery location +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of short circuit in lunar module descent battery +2025-04-03 at 09:40:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:40:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of short circuit in command module umbilical system +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of descent batteries in lunar module +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: C-axis induction potentiometer component failures antenna system +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent stage descent battery 4 quadrant location electrical short circuit +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent battery electrical system location +2025-04-03 at 09:40:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:40:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of electrical short circuit in command module umbilical connectors +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: C-axis induction potentiometer electrically isolated or isolated from the rest of the system +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: current surge location in lunar module descent battery electrical system +2025-04-03 at 09:40:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:40:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:40:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of electrical connectors in command module service module umbilical system +2025-04-03 at 09:40:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:40:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of quadrant 4 in lunar module descent stage +2025-04-03 at 09:40:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:40:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:40:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: specific connectors involved in power supply to command module during transearth phase +2025-04-03 at 09:40:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:40:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of quadrant 4 in lunar module descent stage +2025-04-03 at 09:40:58 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:40:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: _transearth injection maneuver lunar module +2025-04-03 at 09:41:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:41:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:41:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection maneuver lunar module short circuit +2025-04-03 at 09:41:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:41:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:41:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power exhaustion plan +2025-04-03 at 09:41:03 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:41:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:41:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module battery electrical system current surge location +2025-04-03 at 09:41:04 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:41:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:41:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:62 - Student lengths: [626, 522, 676, 861, 307, 2031] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [23, 23, 23, 23, 23, 23] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:64 - Average student length: 837.17 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 23.00 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_correctness:66 - Length ratio: 36.40 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.612 ± 0.323 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.50 ± 3.10 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 3, 5, 0, 3, 10] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 09:41:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...', 'Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...', 'Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...'] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 09:41:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...', 'Result 1:\nFigure 14-6.- Recorded signal strengths during high-gain antenna operation.\n\n...', 'Result 1:\nelectronic box and trigger the antenna logic to produce the scan-limit functi...'] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:41:04 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:41:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the operating temperature of a car engine +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical operating temperature range for a process or system that is often associated with high temperatures and gases? +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: chiffonade gas turbine or jet engine altitude temperature range +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature of gas operation +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperate gas during operating conditions +2025-04-03 at 09:41:10 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module helium manifold pressures and temperatures +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature of the gas in a rocket engine like that used in the Apollo mission? +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gas temperature cabin aircraft NASA +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command module helium temperature Apollo mission +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: engine package temperature during ascent +2025-04-03 at 09:41:14 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank temperature range for spacecraft +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Descent Program gas temperature" or "Gas temperature inside the Apollo Descent Engine +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: impossible to withdraw water after landing command module temperature +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Post Apollo mission command module helium temperatures +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module helium temperatures during ascent +2025-04-03 at 09:41:17 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium temperature for 8 psi/hour rise rate +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Helium and hydrogen gas temperature relationship during lunar descent" or "Lunar Descent Propulsion System gas temperatures during operation +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module temperature near water tank postlanding +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank helium temperature +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium temperature range during ascent +2025-04-03 at 09:41:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium peak temperature during spacecraft descent +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Helium temperature range for special lunar landing timeline +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: low cabin temperature causing water tank issues +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank screening test pressure rise rate +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supercritical helium temperature during ascent +2025-04-03 at 09:41:24 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium temperature at 900 psia +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Relationship between helium temperature and pressure 900 psia" or "Temperature reading equivalent 900 psia helium pressure +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inconsistency in cabin temperature during powered-down and return-flight conditions +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Steady-state rise rate helium tank testing +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supercritical helium temperature +2025-04-03 at 09:41:28 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:41:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank screening test results +2025-04-03 at 09:41:31 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:41:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank operation temperature +2025-04-03 at 09:41:32 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:33 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:41:33 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:41:33 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, False, True, True, False, False] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.50 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.50 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:62 - Student lengths: [390, 322, 435, 1697, 254, 306] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:64 - Average student length: 567.33 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 16.00 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_correctness:66 - Length ratio: 35.46 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.713 ± 0.363 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 5.67 ± 2.69 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [6, 6, 8, 8, 6, 0] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa..."] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa..."] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...'] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", 'Result 1:\nFigure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa..."] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe only anomaly observed in the environmental control system was a reverse l...', 'Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe supercritical helium pressurization system displayed abnormal performance...', 'Result 1:\nThe supercritical helium pressurization system displayed abnormal performance...'] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +2025-04-03 at 09:41:33 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 0.0 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.000 +2025-04-03 at 09:41:33 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:41:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the unidentified signals received during the afternoon of a December 2022 launch of a satellite or rocket?, +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: sporadic signals launch day [rocket launch signals origin] +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the sporadic signals recorded during the launch of the Perseverance Mars rover in February 2021 or any other notable space launch in the recent past? +2025-04-03 at 09:41:39 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:41:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What effects of launch-phase electrical phenomena can cause the afternoon signals detected following a launch of a high-powered rocket?, +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch vehicle exhaust sparks or electromagnetic signals +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What scientific documentation, if any, mentions the origin of the unusual signals recorded by the Shuttle's field-change and sferics detectors on April 12, 1985, or April 12th in NASA missions. +2025-04-03 at 09:41:41 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:41:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of electromagnetic pulse is generated by a launch vehicle's hot exhaust?, +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch electromagnetic interference or electrical charge separation +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What records confirm that the Apollo 13 mission experienced significant electrical charge separation during launch, and what impact did it have on the mission's trajectory and communications? +2025-04-03 at 09:41:44 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:41:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 09:41:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 electromagnetic interference investigation +2025-04-03 at 09:41:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:41:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the electrical charge separation during the Apollo 13 launch impact the mission's orbit, leading to any unusual trajectories or communication issues? +2025-04-03 at 09:41:47 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:41:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo launch electrical charge separation lightning incident +2025-04-03 at 09:41:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:41:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How did the Apollo 13 mission's launches impact command and data transmitting communication systems and was there any significant change in the spacecraft's trajectory due to the discovered phenomena? +2025-04-03 at 09:41:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:41:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrostatic discharge potential of Apollo 13 +2025-04-03 at 09:41:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:41:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were there any deviations in the Apollo 13 spacecraft's trajectory due to the discovered electrical charge separation during launch and its effect on communication? +2025-04-03 at 09:41:51 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:41:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 electrostatic discharge and signal origin +2025-04-03 at 09:41:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:41:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was the launch's charging condition a necessary factor for adjusting the command and data link antenna orientation to increase signal gain during transmission with tracking station Houston, to guarantee the success of some critical phase of the mission? +2025-04-03 at 09:41:54 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:41:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 electrostatic potential and signal origin +2025-04-03 at 09:41:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:41:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the launch's electrical field changes affect the performance of the Apollo 13 spacecraft's power distribution system, despite the evidence suggesting flight computers, Service and power distribution units, Communications equipment, and power sources are all were nominal during mission and for several days after lift-off? +2025-04-03 at 09:41:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:41:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 09:41:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch signals and origin +2025-04-03 at 09:41:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:41:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were the launch's electrical discharge differences between launch lifetime expected to have been contributed by amount of electrical power in the system generated as part of temperature rise as the rocket cooled down after engine cut off, by review of required insignia system performance? +2025-04-03 at 09:41:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 09:41:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:42:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch signals and training +2025-04-03 at 09:42:00 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:42:00 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:42:00 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:42:00 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, False, True, True, True] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:56 - Average correctness: 0.83 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.37 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1078, 577, 1956, 1581, 314, 28] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:64 - Average student length: 922.33 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 25.00 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_correctness:66 - Length ratio: 36.89 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.396 ± 0.426 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 4.00 ± 4.28 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 1/6 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [0, 5, 10, 9, 0, 0] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 9: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 09:42:00 | WARNING | src.rewards:reward_em_chunk:296 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 3.0 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 0.500 +2025-04-03 at 09:42:00 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.500 +2025-04-03 at 09:42:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "What are the effects of in-flight chlorination on bacterial populations in the air?" +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did in-flight chlorination of a welding respirator by NASA experiment prove to be effective in eliminating bacteria? +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Fluorocarbon bacteria survival rate on aircraft surfaces after chlorination +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the effects of in-flight chlorination on bacterial populations in birds? +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the outcome of bacterial populations in aircraft toilets after chlorination? +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the growth of S. aureus and other bacteria in an aircraft cabin after chlorination? +2025-04-03 at 09:42:07 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Effectiveness of chlorination on Apollo 17 command module water system" +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were in-flight bacterial growths or other issues caused by the closed fuel valve or miswiring in the lunar module during or after the propellant dump and purge? +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of chlorination on bacterial growth in closed systems +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the outcome of the bacterial population after crew members consumed naturally chlorinated water from unchlorinated water supply after Skylab 3 inflight chlorination? +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the effect on bacterial population after chlorination in the Apollo 11 space toilet? +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effect of in-flight chlorination on bacterial population in Apollo 13 crew's command module water system +2025-04-03 at 09:42:11 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Did inflight chlorination of lunar-modular lunar lander's system affect its performance, reuse or any issues?" +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were in-flight chlorination experiments specific to the lunar module's water system designed to eliminate bacteria, and were chlorination procedures followed strictly and successfully for the entire duration of the mission? +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of in-flight chlorination on bacterial population +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Skylab 3 crew bacterial population count after in-flight chlorination success +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effectiveness of in-flight chlorination on reducing bacterial populations in the Apollo space toilet +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effect of water usage rate on bacterial population during extended stay in reduced crew pressure environments with limited access to resupply and decreasing water volumes +2025-04-03 at 09:42:16 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:42:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module water usage discovery volumes of water consumption per hr. Were there water usage issues and any incrment for cads of water?" +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the outcome of bacterial growth after inflight chlorination in the lunar module water system during the Apollo mission? +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effectiveness of in-flight chlorination in reducing bacterial populations +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Skylab 3 crew water usage and bacterial growth after inflight chlorination +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - �� Search Query: effectiveness of in-flight bacterial reduction in the Apollo 11 space toilet after chlorination +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effect of exhaustion of potable water supply and lack of backup system on bacterial population control in the Apollo 13 command module +2025-04-03 at 09:42:21 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much water (lunar moon) did command module utilise +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the lunar module's water conservation system and crew water usage affect the outcome of bacterial growth during the Apollo mission? +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effect of in-flight chlorination on bacterial population in absence of subsequent chlorination +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Drinking water usage and bacterial growth in the Skylab 3 crew after inflight chlorination +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effect of in-flight chlorination on total bacterial count in Apollo 11 command module water system +2025-04-03 at 09:42:26 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the command module water usage insulation a bacterial growth in the water system +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Were the water usage rates in the lunar module sufficient to reduce bacterial growth, or did the limited water supply necessitate more frequent chlorination? +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: persistent effects of in-flight chlorination on bacterial populations +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: determinants of bacterial growth during Skylab 3 crew water usage from lunar module +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lack of available information on bacterial population after chlorination in the Apollo 11 command module +2025-04-03 at 09:42:30 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:42:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "What effect did cryogenic must vent finitude in oxygen tank impact command module operations +2025-04-03 at 09:42:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:42:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Comparison of bacterial counts in Skylab 3 command module and lunar module water tanks after inflight chlorination +2025-04-03 at 09:42:35 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:42:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Skylab 3 command module water bacterial count before inflight chlorination +2025-04-03 at 09:42:36 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 09:42:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 09:42:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 09:42:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:55 - Correctness metrics: [True, True, True, True, True, True] +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:56 - Average correctness: 1.00 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:57 - Standard deviation: 0.00 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:62 - Student lengths: [1692, 1210, 681, 747, 620, 642] +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:63 - Teacher lengths: [23, 23, 23, 23, 23, 23] +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:64 - Average student length: 932.00 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:65 - Average teacher length: 23.00 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_correctness:66 - Length ratio: 40.52 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_format:164 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_retry:245 - Retry behavior rewards: 0.792 ± 0.224 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_retry:246 - Search tags per completion: 7.50 ± 3.04 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_retry:247 - Violations (>1 search per message): 2/6 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_retry:248 - Search counts distribution: [14, 6, 6, 8, 6, 5] +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.· To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 7: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 8: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:288 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 3: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:290 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:306 - Chunk Query Rewards Summary: +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:307 - Total prompts: 6 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:308 - Correct matches: 6.0 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:309 - Average reward: 1.000 +2025-04-03 at 09:42:38 | INFO | src.rewards:reward_em_chunk:310 - Reward std: 0.000 +2025-04-03 at 09:42:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propulsion auxiliary propellant gaging system failure cause +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What caused failure of service propulsion auxiliary propellant gaging system. +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propellant gaging system failure Lockheed Orion spacecraft +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: failure of service propulsion auxiliary propellant gaging system +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service propulsion auxiliary propellant gaging system NASA failure +2025-04-03 at 09:42:46 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:42:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel leakage into point sensor module propellant gaging system failure +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reason for failure of service propulsion auxiliary propellant gaging system due to fuel leakage. +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 propellant gaging system failure +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reason for fuel leakage in service propulsion auxiliary propellant gaging system +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 service propulsion auxiliary propellant gaging system failure +2025-04-03 at 09:42:49 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:42:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel leak point sensor module propellant gaging system +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apex cover jettison system issue propellant gaging system failure. +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 propulsion system anomaly reference l +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: boil-off of O-ring backed Teflon ring in gusset-4 breech assembly +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 systems failure guidance propulsion auxiliary propellant gaging system +2025-04-03 at 09:42:53 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 09:42:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service propulsion auxiliary propellant gaging system electrical circuit interrupter location +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 abort profile section 3 +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: excessive O-ring compression during valve assembly +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 propellant management system failure +2025-04-03 at 09:42:56 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 09:42:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: left-hand electrical circuit interrupter Service Propulsion Auxiliary Propellant Gaging System +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 abort profile Apollo 11 service propulsion system +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reason for propellant gaging system failure +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 abort command module propulsion system +2025-04-03 at 09:42:59 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 09:42:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: O-ring failure in electrical circuit interrupts Service Propulsion Auxiliary Propellant Gaging System +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 abort profile Apollo 11 cryogenic oxygen supply +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service propulsion auxiliary propellant gaging system design or manufacturing issue +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zer， lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system auxiliary propellant gaging system failure +2025-04-03 at 09:43:02 | INFO | src.agent:run_tool_calls:118 - ℹ️ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 09:43:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts